NPrize nprize https://lcas.otaski.org/index.php?title=Main_Page MediaWiki 1.42.3 first-letter Media Special Talk User User talk NPrize NPrize talk File File talk MediaWiki MediaWiki talk Template Template talk Help Help talk Category Category talk 0 1 1 2010-01-04T15:33:36Z MediaWiki default 0 wikitext text/x-wiki <big>'''MediaWiki has been successfully installed.'''</big> Consult the [http://meta.wikimedia.org/wiki/Help:Contents User's Guide] for information on using the wiki software. == Getting started == * [http://www.mediawiki.org/wiki/Manual:Configuration_settings Configuration settings list] * [http://www.mediawiki.org/wiki/Manual:FAQ MediaWiki FAQ] * [http://lists.wikimedia.org/mailman/listinfo/mediawiki-announce MediaWiki release mailing list] 08fc850f2898611c250d639e30f69532b5a016f8 2 1 2010-01-04T15:47:36Z Vincent 1 wikitext text/x-wiki <big>'''N-Prize reflexion wiki'''</big> This web site aims to gather free information on how to challenge the N-Prize competition. == Getting started == * [http://www.mediawiki.org/wiki/Manual:Configuration_settings Configuration settings list] * [http://www.mediawiki.org/wiki/Manual:FAQ MediaWiki FAQ] * [http://lists.wikimedia.org/mailman/listinfo/mediawiki-announce MediaWiki release mailing list] 97903fb0cfe45a0c6a74a176bdfc68fcf5f2029a 4 2 2010-01-04T16:07:17Z Vincent 1 wikitext text/x-wiki <big>'''N-Prize reflexion wiki'''</big> This web site aims to gather free information on how to challenge the N-Prize competition. It is not an official web site for the N-Prize. The official web site is here: http://www.n-prize.com/ 33b8b4eeb2fa3a284f70842739213886fe71cd87 8 4 2010-01-04T16:24:25Z Vincent 1 Import from html wikitext text/x-wiki <big>'''N-Prize reflexion wiki'''</big> This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the N-Prize. The official web site is here: http://www.n-prize.com/ . I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for two things: go fast and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds. Besides altitude, speed is the most important factor when orbiting an object. Without it, satellites would fall down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, while the final direction of the flight is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection. That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbo-propeller is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design is also part of the improvement. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to orbit a payload up to 1,000 lbs (450 kg), and is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft because I am not a millionaire. A rocket would still be used for air-to-orbit link because nothing else is able to achieve speeds around 9 km/s. Some specific technologies can be used to improve efficiency, we'll see them below in the rocket section. Several teams are working on using Helium balloons to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and less expensive in the overall, but balloons are not reusable and suffer from imprecise position localization due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbo-propeller? Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece. ==The rocket== ===Fuel=== represents the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Trajectory=== has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 238eabaaf8d3deb41e40763861257c6fc23dfcab 9 8 2010-01-04T17:39:41Z Vincent 1 /* How to escape from Earth? */ wikitext text/x-wiki <big>'''N-Prize reflexion wiki'''</big> This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the N-Prize. The official web site is here: http://www.n-prize.com/ . I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for two things: go fast and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds. Besides altitude, speed is the most important factor when orbiting an object. Without it, satellites would fall down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, while the final direction of the flight is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection. That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbo-propeller is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design is also part of the improvement. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to orbit a payload up to 1,000 lbs (450 kg), and is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft because I am not a millionaire. A rocket would still be used for air-to-orbit link because nothing else is able to achieve speeds around 9 km/s. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket]] section. Several teams are working on using Helium balloons to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and less expensive in the overall, but balloons are not reusable and suffer from imprecise position localization due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbo-propeller? Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece. ==The rocket== ===Fuel=== represents the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Trajectory=== has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 062e8db51c627d8029b8d90889be770c09948192 10 9 2010-01-04T17:40:34Z Vincent 1 /* How to escape from Earth? */ wikitext text/x-wiki <big>'''N-Prize reflexion wiki'''</big> This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the N-Prize. The official web site is here: http://www.n-prize.com/ . I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for two things: go fast and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds. Besides altitude, speed is the most important factor when orbiting an object. Without it, satellites would fall down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, while the final direction of the flight is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection. That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbo-propeller is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design is also part of the improvement. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to orbit a payload up to 1,000 lbs (450 kg), and is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft because I am not a millionaire. A rocket would still be used for air-to-orbit link because nothing else is able to achieve speeds around 9 km/s. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium balloons to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and less expensive in the overall, but balloons are not reusable and suffer from imprecise position localization due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbo-propeller? Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece. ==The rocket== ===Fuel=== represents the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Trajectory=== has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 42ec75126d212844b42a0c06bc5e0cdaeada5683 11 10 2010-01-04T19:54:20Z Vincent 1 wikitext text/x-wiki <big>'''N-Prize reflexion wiki'''</big> This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the N-Prize. The official web site is here: http://www.n-prize.com/ . I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for two things: go fast and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds. Besides altitude, speed is the most important factor when orbiting an object. Without it, satellites would fall down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, while the final direction of the flight is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection. That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbo-propeller is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design is also part of the improvement. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to orbit a payload up to 1,000 lbs (450 kg), and is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft because I am not a millionaire. A rocket would still be used for air-to-orbit link because nothing else is able to achieve speeds around 9 km/s. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium balloons to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and less expensive in the overall, but balloons are not reusable and suffer from imprecise position localization due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbo-propeller? ===Staging=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 4f67907e6f39946b7b658ca14d536daa8ea70cb6 16 11 2010-01-05T02:23:31Z Vincent 1 add N-Prize internal link wikitext text/x-wiki <big>'''N-Prize reflexion wiki'''</big> This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for two things: go fast and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds. Besides altitude, speed is the most important factor when orbiting an object. Without it, satellites would fall down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, while the final direction of the flight is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection. That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbo-propeller is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design is also part of the improvement. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to orbit a payload up to 1,000 lbs (450 kg), and is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft because I am not a millionaire. A rocket would still be used for air-to-orbit link because nothing else is able to achieve speeds around 9 km/s. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium balloons to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and less expensive in the overall, but balloons are not reusable and suffer from imprecise position localization due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbo-propeller? ===Staging=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... f2193483ebaeaf8579a4f2356f981020c682189a 18 16 2010-01-05T02:41:24Z Vincent 1 wikitext text/x-wiki <big>'''N-Prize reflexion wiki'''</big> This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for two things: go fast and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds. Besides altitude, speed is the most important factor when orbiting an object. Without it, satellites would fall down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, while the final direction of the flight is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection. That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbo-propeller is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design is also part of the improvement. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve speeds around 9 km/s. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium balloons to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and less expensive in the overall, but balloons are not reusable and suffer from imprecise position localization due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbo-propeller? ===Staging=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxydizer=== Using LOX has tons of drawbacks because of cryogenics. But are there any other oxydizers that can be used? ===Engine=== Aerospike engines should be seriously considered. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 8c79e0f1ffe931e792c001e271419747829c9775 20 18 2010-01-22T11:33:52Z Vincent 1 embedded computer link wikitext text/x-wiki <big>'''N-Prize reflexion wiki'''</big> This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for two things: go fast and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds. Besides altitude, speed is the most important factor when orbiting an object. Without it, satellites would fall down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, while the final direction of the flight is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection. That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbo-propeller is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design is also part of the improvement. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve speeds around 9 km/s. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium balloons to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and less expensive in the overall, but balloons are not reusable and suffer from imprecise position localization due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbo-propeller? ===Staging=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxydizer=== Using LOX has tons of drawbacks because of cryogenics. But are there any other oxydizers that can be used? ===Engine=== Aerospike engines should be seriously considered. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 49c05214383c3522ded59b1c83502baff45aa19a 22 20 2010-01-22T17:06:56Z Vincent 1 /* Engine */ aerospike links wikitext text/x-wiki <big>'''N-Prize reflexion wiki'''</big> This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for two things: go fast and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds. Besides altitude, speed is the most important factor when orbiting an object. Without it, satellites would fall down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, while the final direction of the flight is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection. That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbo-propeller is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design is also part of the improvement. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve speeds around 9 km/s. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium balloons to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and less expensive in the overall, but balloons are not reusable and suffer from imprecise position localization due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbo-propeller? ===Staging=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxydizer=== Using LOX has tons of drawbacks because of cryogenics. But are there any other oxydizers that can be used? ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... f84c23b074549669478a99b238c76ce649d62628 23 22 2010-01-22T17:36:01Z Vincent 1 /* How to escape from Earth? */ thrust wikitext text/x-wiki <big>'''N-Prize reflexion wiki'''</big> This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds. Besides altitude, speed is the most important factor when orbiting an object. Without it, satellites would fall down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, while the final direction of the flight is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection. That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbo-propeller is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design is also part of the improvement. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve speeds around 9 km/s. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium balloons to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and less expensive in the overall, but balloons are not reusable and suffer from imprecise position localization due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbo-propeller? ===Staging=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxydizer=== Using LOX has tons of drawbacks because of cryogenics. But are there any other oxydizers that can be used? ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... e8498e2739ea7c996fdc25bedda9642463cb153e 25 23 2010-01-22T22:56:37Z Vincent 1 reviewing and more explanations wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbo-propeller is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable and suffer from imprecise trajectory due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbo-propeller? ===Staging=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxydizer=== Using LOX has tons of drawbacks because of cryogenics. But are there any other oxydizers that can be used? ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 1d88861a3c98dbb5a51b2ffb24fae6a192fea2e2 27 25 2010-01-22T23:49:51Z Vincent 1 /* How to escape from Earth? */ picture wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. <center>[[Image:Rocket_trajectory.png]]</center> That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbo-propeller is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable and suffer from imprecise trajectory due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbo-propeller? ===Staging=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxydizer=== Using LOX has tons of drawbacks because of cryogenics. But are there any other oxydizers that can be used? ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 2ed1444fb14878b26a44bde4ecc0229fa60b0377 30 27 2010-01-22T23:59:56Z Vincent 1 /* How to escape from Earth? */ change centering image attributes wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbo-propeller is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable and suffer from imprecise trajectory due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbo-propeller? ===Staging=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxydizer=== Using LOX has tons of drawbacks because of cryogenics. But are there any other oxydizers that can be used? ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... d684de802b640b3431c32a6801176c4d53235bc0 31 30 2010-01-25T15:41:19Z Vincent 1 /* Oxydizer */ wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbo-propeller is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable and suffer from imprecise trajectory due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbo-propeller? ===Staging=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxydizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has tons of drawbacks because of the need for cryogenics storage, manipulation, and engine design, that make it quite expensive and too much complicated. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 83d852a3759907f0308b06006d02337d5592269e 32 31 2010-01-25T15:52:50Z Vincent 1 propeller->fan wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable and suffer from imprecise trajectory due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, flights at this altitude at 20km/h. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Fuel or electricity? Kerozene or alcohol? How to build a £100 turbofan? ===Staging=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxydizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has tons of drawbacks because of the need for cryogenics storage, manipulation, and engine design, that make it quite expensive and too much complicated. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 0f8accb90d5b5d359eb1268dee7a73a0370a4461 34 32 2010-01-25T18:15:56Z Vincent 1 /* The aircraft */ wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable and suffer from imprecise trajectory due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but crusing at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerozene, alcohol, E85? Anyway, the main issue with the aircraft is : ''how to build a £100 turbofan?'' ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxydizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has tons of drawbacks because of the need for cryogenics storage, manipulation, and engine design, that make it quite expensive and too much complicated. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 4dc576fb7f1690176704fcd0c65c7a688c3a5672 35 34 2010-01-25T18:57:57Z Vincent 1 /* Oxydizer */ wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxydizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable and suffer from imprecise trajectory due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but crusing at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerozene, alcohol, E85? Anyway, the main issue with the aircraft is : ''how to build a £100 turbofan?'' ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxydizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has tons of drawbacks because of the need for cryogenics storage, manipulation, and engine design, that make it quite expensive and too much complicated. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 19ca709ec7357257af204bcbed4791ab42f3c442 36 35 2010-01-25T19:07:14Z Vincent 1 oxydizer -> oxidizer wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching, and I found out that Orbital already has developped an air-to-orbit vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 70km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable and suffer from imprecise trajectory due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but crusing at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerozene, alcohol, E85? Anyway, the main issue with the aircraft is : ''how to build a £100 turbofan?'' ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has tons of drawbacks because of the need for cryogenics storage, manipulation, and engine design, that make it quite expensive and too much complicated. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... b598d07323c16abe05eb94db27ceb944e756a1fe 40 36 2010-01-26T16:37:24Z Vincent 1 wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable and suffer from imprecise trajectory due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but crusing at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerozene, alcohol, E85? Anyway, the main issue with the aircraft is : ''how to build a £100 turbofan?'' ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has tons of drawbacks because of the need for cryogenics storage, manipulation, and engine design, that make it quite expensive and too much complicated. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 6a557db12aa630e427fc7630fc0b7d82b3bd3ba9 42 40 2010-01-26T19:20:50Z Vincent 1 /* The aircraft */ turbofan link wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable and suffer from imprecise trajectory due to winds. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but crusing at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerozene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has tons of drawbacks because of the need for cryogenics storage, manipulation, and engine design, that make it quite expensive and too much complicated. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... a06d88149acf6bd18203da4319e1eba0d0fdfcaa 8 2 3 2010-01-04T16:03:40Z Vincent 1 New page: * navigation ** mainpage|mainpage ** recentchanges-url|recentchanges ** randompage-url|randompage wikitext text/x-wiki * navigation ** mainpage|mainpage ** recentchanges-url|recentchanges ** randompage-url|randompage cf0a30d123e7eba0b28c6708e3ef8348e6be891e 46 3 2010-01-29T09:35:50Z Vincent 1 adding links to computer and turbofan wikitext text/x-wiki * navigation ** mainpage|mainpage *** EmbeddedRocketComputer|Computer *** Build_a_cheap_turbofan|Plane engine ** recentchanges-url|recentchanges ** randompage-url|randompage 2d766bdca882c7ab0926ed42f7bef9dca3b8c0d3 47 46 2010-01-29T09:40:03Z Vincent 1 split with new section wikitext text/x-wiki * navigation ** mainpage|mainpage ** recentchanges-url|recentchanges ** randompage-url|randompage * sections ** EmbeddedRocketComputer|Computer ** Build_a_cheap_turbofan|Plane engine bcb5bfe99dc2fbc614f766e9fa4527f38e44d8b2 12 3 5 2010-01-04T16:08:01Z Vincent 1 New page: == Getting started == * [http://www.mediawiki.org/wiki/Manual:Configuration_settings Configuration settings list] * [http://www.mediawiki.org/wiki/Manual:FAQ MediaWiki FAQ] * [http://list... wikitext text/x-wiki == Getting started == * [http://www.mediawiki.org/wiki/Manual:Configuration_settings Configuration settings list] * [http://www.mediawiki.org/wiki/Manual:FAQ MediaWiki FAQ] * [http://lists.wikimedia.org/mailman/listinfo/mediawiki-announce MediaWiki release mailing list] 9a1c8b7d03b1110ae87fb98761303bb648d4f224 12 4 6 2010-01-04T16:16:56Z Vincent 1 New page: {{PD Help Page}} You can format your text using wiki markup. This consists of normal characters like asterisks, single quotes or equation marks which have a special function in the wiki, s... wikitext text/x-wiki {{PD Help Page}} You can format your text using wiki markup. This consists of normal characters like asterisks, single quotes or equation marks which have a special function in the wiki, sometimes depending on their position. For example, to format a word in ''italic'', you include it in two single quotes like <code><nowiki>''this''</nowiki></code>. == Text formatting markup == {| class="wikitable" ! Description !! You type !! You get |- ! colspan="3" style="background:#ABE" | character (inline) formatting – ''applies anywhere'' |- |Italic text | <code><nowiki>''italic''</nowiki></code> |''italic'' |- | Bold text | <code><nowiki>'''bold'''</nowiki></code> |'''bold''' |- | Bold and italic | <code><nowiki>'''''bold & italic'''''</nowiki></code> |'''''bold & italic''''' |- |Escape wiki markup | <code><nowiki><nowiki>no ''markup''&lt;/nowiki></nowiki></code> |<nowiki>no ''markup''</nowiki> |- ! colspan="3" style="background:#ABE" | section formatting – ''only at the beginning of the line'' |- |Headings of different levels | <pre>=level 1= ==level 2== ===level 3=== ====level 4==== =====level 5===== ======level 6======</pre> An article with 4 or more headings automatically creates a [[wikipedia:Wikipedia:Section#Table of contents (TOC)|table of contents]]. |<!-- hack to prevent TOC viewing for h1 - h6 elements: their style is hardcopied here --> <div style="font-size: 188%; margin: 0; padding-top: .5em; padding-bottom: .17em; border-bottom: 1px solid #aaa">Level 1</div> <div style="font-size: 150%; margin: 0; padding-top: .5em; padding-bottom: .17em; border-bottom: 1px solid #aaa">Level 2</div> <div style="font-size: 132%; font-weight: bold">Level 3</div><!-- --><b>Level 4</b><!-- --><div style="font-size: 86%; font-weight: bold">Level 5</div><!-- --><b style="font-size: 80%">Level 6</b> |- |Horizontal rule | <code>----</code> | ---- |- |Bullet list | <pre> * one * two * three ** three point one ** three point two </pre> Inserting a blank line will end the first list and start another. | * one * two * three ** three point one ** three point two |- |Numbered list | <pre> # one # two<br />spanning more lines<br />doesn't break numbering # three ## three point one ## three point two </pre> | # one # two<br />spanning more lines<br />doesn't break numbering # three ## three point one ## three point two |- |Definition list |<pre> ;item 1 : definition 1 ;item 2 : definition 2-1 : definition 2-2 </pre> | ;item 1 : definition 1 ;item 2 : definition 2-1 : definition 2-2 |- | Adopting definition list to indent text | <pre>: Single indent :: Double indent ::::: Multiple indent</pre> This workaround may be controversial from the viewpoint of accessibility. | : Single indent :: Double indent ::::: Multiple indent |- | Mixture of different types of list | <pre> # one # two #* two point one #* two point two # three #; three item one #: three def one # four #: four def one #: this rather looks like the continuation of # four #: and thus often used instead of <br /> # five ## five sub 1 ### five sub 1 sub 1 ## five sub 2 ;item 1 :* definition 1-1 :* definition 1-2 : ;item 2 :# definition 2-1 :# definition 2-2 </pre> The usage of <code>#:</code> and <code>*:</code> for breaking a line within an item may also be controversial. | # one # two #* two point one #* two point two # three #; three item one #: three def one # four #: four def one #: this rather looks like the continuation of <code># four</code> #: often used instead of <code>&lt;br /></code> # five ## five sub 1 ### five sub 1 sub 1 ## five sub 2 ;item 1 :* definition 1-1 :* definition 1-2 : ;item 2 :# definition 2-1 :# definition 2-2 |- |- |Preformatted text | <pre> preformatted text is done with a '''space''' at the ''beginning'' of the line </pre> This way of preformatting only applies to section formatting, and character formatting markups are still effective. | preformatted text is done with a '''space''' at the ''beginning'' of the line |} == Paragraphs == MediaWiki ignores normal line breaks. To start a new paragraph, leave an empty line. You can force a line break within a paragraph with the HTML tags <code>&lt;br /></code>. == HTML == Some [[wikipedia:HTML|HTML]] tags are allowed in MediaWiki, for example <code>&lt;code></code>, <code>&lt;div></code>, <code><nowiki><span></nowiki></code> and <code><nowiki><font></nowiki></code>. These apply anywhere you insert them. {| class="wikitable" !Description !You type !You get |- | Strikethrough | <code>&lt;del>Strikethrough&lt;/del></code> or <code>&lt;s>Strikethrough&lt;/s></code> |<del>Strikethrough</del> |- style="white-space:nowrap" |Fixed width text | <code>&lt;tt>Fixed width text&lt;/tt></code> or <code>&lt;code>source code&lt;/code></code> | <tt>Fixed width text</tt> |- |Blockquotes | <code> text text text text text text text text text text text text text text text text text text text text text text text text &lt;blockquote> quote quote quote quote quote quote &lt;/blockquote> text text text text text text text text text text text text </code> | text text text text text text text text text text text text text text text text text text text text text text text text <blockquote> quote quote quote quote quote quote </blockquote> text text text text text text text text text text text text |- |Comment | <code>&lt;!-- This is a comment --></code> Text can only be viewed in the edit window. | <!-- This is a real invisible comment --> |- |Completely preformatted text | <code><nowiki><pre>this way, all markups are '''ignored'''</pre></nowiki></code> | <pre> this way, all markups are '''ignored'''.</pre> |- |'''Customised''' preformatted text | <code><nowiki><pre style="CSS text">this way, all markups are '''ignored''' and formatted with a CSS text</pre></nowiki></code> | <pre style="white-space:pre-wrap;white-space:-moz-pre-wrap;white-space:-pre-wrap;white-space:-o-pre-wrap;word-wrap:break-word;overflow:auto;"> this way for instance, all '''ignored''' markups take into account the navigator size, by automatically adding some carriage returns dynamically to it.</pre> |} {{Languages}} [[Category:Help|Formatting]] cbed50551017770dc06895041f73562a50509ba6 7 6 2010-01-04T16:23:11Z Vincent 1 links wikitext text/x-wiki You can format your text using wiki markup. This consists of normal characters like asterisks, single quotes or equation marks which have a special function in the wiki, sometimes depending on their position. For example, to format a word in ''italic'', you include it in two single quotes like <code><nowiki>''this''</nowiki></code>. == Text formatting markup == {| class="wikitable" ! Description !! You type !! You get |- ! colspan="3" style="background:#ABE" | character (inline) formatting – ''applies anywhere'' |- |Italic text | <code><nowiki>''italic''</nowiki></code> |''italic'' |- | Bold text | <code><nowiki>'''bold'''</nowiki></code> |'''bold''' |- | Bold and italic | <code><nowiki>'''''bold & italic'''''</nowiki></code> |'''''bold & italic''''' |- |Escape wiki markup | <code><nowiki><nowiki>no ''markup''&lt;/nowiki></nowiki></code> |<nowiki>no ''markup''</nowiki> |- ! colspan="3" style="background:#ABE" | section formatting – ''only at the beginning of the line'' |- |Headings of different levels | <pre>=level 1= ==level 2== ===level 3=== ====level 4==== =====level 5===== ======level 6======</pre> An article with 4 or more headings automatically creates a [[wikipedia:Wikipedia:Section#Table of contents (TOC)|table of contents]]. |<!-- hack to prevent TOC viewing for h1 - h6 elements: their style is hardcopied here --> <div style="font-size: 188%; margin: 0; padding-top: .5em; padding-bottom: .17em; border-bottom: 1px solid #aaa">Level 1</div> <div style="font-size: 150%; margin: 0; padding-top: .5em; padding-bottom: .17em; border-bottom: 1px solid #aaa">Level 2</div> <div style="font-size: 132%; font-weight: bold">Level 3</div><!-- --><b>Level 4</b><!-- --><div style="font-size: 86%; font-weight: bold">Level 5</div><!-- --><b style="font-size: 80%">Level 6</b> |- |Horizontal rule | <code>----</code> | ---- |- |Bullet list | <pre> * one * two * three ** three point one ** three point two </pre> Inserting a blank line will end the first list and start another. | * one * two * three ** three point one ** three point two |- |Numbered list | <pre> # one # two<br />spanning more lines<br />doesn't break numbering # three ## three point one ## three point two </pre> | # one # two<br />spanning more lines<br />doesn't break numbering # three ## three point one ## three point two |- |Definition list |<pre> ;item 1 : definition 1 ;item 2 : definition 2-1 : definition 2-2 </pre> | ;item 1 : definition 1 ;item 2 : definition 2-1 : definition 2-2 |- | Adopting definition list to indent text | <pre>: Single indent :: Double indent ::::: Multiple indent</pre> This workaround may be controversial from the viewpoint of accessibility. | : Single indent :: Double indent ::::: Multiple indent |- | Mixture of different types of list | <pre> # one # two #* two point one #* two point two # three #; three item one #: three def one # four #: four def one #: this rather looks like the continuation of # four #: and thus often used instead of <br /> # five ## five sub 1 ### five sub 1 sub 1 ## five sub 2 ;item 1 :* definition 1-1 :* definition 1-2 : ;item 2 :# definition 2-1 :# definition 2-2 </pre> The usage of <code>#:</code> and <code>*:</code> for breaking a line within an item may also be controversial. | # one # two #* two point one #* two point two # three #; three item one #: three def one # four #: four def one #: this rather looks like the continuation of <code># four</code> #: often used instead of <code>&lt;br /></code> # five ## five sub 1 ### five sub 1 sub 1 ## five sub 2 ;item 1 :* definition 1-1 :* definition 1-2 : ;item 2 :# definition 2-1 :# definition 2-2 |- |- |Preformatted text | <pre> preformatted text is done with a '''space''' at the ''beginning'' of the line </pre> This way of preformatting only applies to section formatting, and character formatting markups are still effective. | preformatted text is done with a '''space''' at the ''beginning'' of the line |} == Paragraphs == MediaWiki ignores normal line breaks. To start a new paragraph, leave an empty line. You can force a line break within a paragraph with the HTML tags <code>&lt;br /></code>. == HTML == Some [[wikipedia:HTML|HTML]] tags are allowed in MediaWiki, for example <code>&lt;code></code>, <code>&lt;div></code>, <code><nowiki><span></nowiki></code> and <code><nowiki><font></nowiki></code>. These apply anywhere you insert them. {| class="wikitable" !Description !You type !You get |- | Strikethrough | <code>&lt;del>Strikethrough&lt;/del></code> or <code>&lt;s>Strikethrough&lt;/s></code> |<del>Strikethrough</del> |- style="white-space:nowrap" |Fixed width text | <code>&lt;tt>Fixed width text&lt;/tt></code> or <code>&lt;code>source code&lt;/code></code> | <tt>Fixed width text</tt> |- |Blockquotes | <code> text text text text text text text text text text text text text text text text text text text text text text text text &lt;blockquote> quote quote quote quote quote quote &lt;/blockquote> text text text text text text text text text text text text </code> | text text text text text text text text text text text text text text text text text text text text text text text text <blockquote> quote quote quote quote quote quote </blockquote> text text text text text text text text text text text text |- |Comment | <code>&lt;!-- This is a comment --></code> Text can only be viewed in the edit window. | <!-- This is a real invisible comment --> |- |Completely preformatted text | <code><nowiki><pre>this way, all markups are '''ignored'''</pre></nowiki></code> | <pre> this way, all markups are '''ignored'''.</pre> |- |'''Customised''' preformatted text | <code><nowiki><pre style="CSS text">this way, all markups are '''ignored''' and formatted with a CSS text</pre></nowiki></code> | <pre style="white-space:pre-wrap;white-space:-moz-pre-wrap;white-space:-pre-wrap;white-space:-o-pre-wrap;word-wrap:break-word;overflow:auto;"> this way for instance, all '''ignored''' markups take into account the navigator size, by automatically adding some carriage returns dynamically to it.</pre> |} ---- == Internal links == To add an internal link, enclose the name of the page you want to link to in double square brackets. When you save the page, you'll see the new link pointing to your page. If the page exists already it is displayed in blue, if it does not, in red. Selflinks to the current page are not transformed in URLs but displayed in bold. (If you really want to link to the current page, use an anchor (see below), or <nowiki>[[#top|current page]]</nowiki> which always links to the top.) The first letter of the target page is automatically capitalized, unless otherwise set by the admins, and spaces are represented as underscores (typing an underscore in the link will have a similar effect as typing a space, but is not recommended, since the underscore will also be shown in the text). {| border="1" class="wikitable" !Description !You type !You get |- |Internal link |<pre>[[Main Page]]</pre> |[[Main Page]] |- |Piped link |<pre>[[Main Page|different text]]</pre> |[[Main Page|different text]] |- |Shortened sort-of-piped link |<pre>[[Internationalisation]]s</pre> |[[Internationalisation]]s |- |Redirect |<pre>#REDIRECT [[Main Page]]</pre> See also [[Help:Redirects]] | &rarr; [[Main Page]] |- |Internal link to an anchor |<pre>[[#See also]]</pre> [[Help:Formatting|Section headings]] and the [[#top|top]] of the page are automatically anchored. |[[#See also]] |- |Internal link to an anchor at another page |<pre>[[Help:Images#See also]]</pre> |[[Help:Images#See also]] |- |Internal link to the current page's talk page |<pre>[[{{TALKPAGENAME}}|Discussion]]</pre> See also [[Help:Magic_words#Page_names]] |[[{{TALKPAGENAME}}|Discussion]] |- |Internal link to a subpage |<pre>[[/example]]</pre> Shortcut for <nowiki>[[Help:Links/example]]</nowiki>. See also [[Help:Subpages]]. |[[/example]] |- |Internal link to a category page |<pre>[[:Category:Help]]</pre> See also [[Help:Categories]] |[[:Category:Help]] |- |Internal link to an image or a file of other types |<pre>[[media:example.jpg]] [[media:example.pdf]]</pre> See also [[Help:Images]] |[[media:example.jpg]] [[media:example.pdf]] |} == External links == {| border="1" class="wikitable" !style="width:15%"|Description !style="width:45%"|You type !style="width:40%"|You get |- |External link |<pre>http://mediawiki.org</pre> |http://mediawiki.org |- |External link with different label |<pre>[http://mediawiki.org MediaWiki]</pre> |[http://mediawiki.org MediaWiki] |- |External link numbered |<pre>[http://mediawiki.org]</pre> |[http://mediawiki.org] |- |External link icons |<pre> [http://en.wikipedia.org/wiki/.avi video] [http://en.wikipedia.org/wiki/.wav sound] [http://en.wikipedia.org/wiki/.pdf document] </pre> External link icons may differ depending on the file type of the target. | [http://en.wikipedia.org/wiki/.avi video] [http://en.wikipedia.org/wiki/.wav sound] [http://en.wikipedia.org/wiki/.pdf document] |- |External link to the same host |<pre>[http://{{SERVERNAME}}/pagename]</pre> |[http://{{SERVERNAME}}/pagename] |- |External link to other host passing the pagename |<pre>[http://google.com/search?q={{PAGENAMEE}}]</pre> |[http://google.com/search?q={{PAGENAMEE}}] |- |Mailto link |<pre>[mailto:info@example.org email me]</pre> |[mailto:info@example.org email me] |- |Mailto named with subject line and body |<pre>[mailto:info@example.org?Subject=URL%20Encoded%20Subject&body=Body%20Text info]</pre> |[mailto:info@example.org?Subject=URL%20Encoded%20Subject&body=Body%20Text info] |} {{admin tip|tip= Which protocols (like http:) are allowed for links is controlled by the {{mediawiki|Manual:$wgUrlProtocols|$wgUrlProtocols}}<!--Should these admin tips even be here? This is supposed to be end user help is it not? --> setting. }} {{admin tip|tip= To remove the “external link icons“ from next to each of the external links, add the following to the page located at <code>MediaWiki:Monobook.css</code> on your wiki. <source lang="css"> #bodyContent a.external, #bodyContent a[href ^="gopher://"] { background: none; padding-right: 0; } </source> }} === How to avoid auto-links === By default, when you write a URL as is, it will be transformed to an external link. To avoid that effect, put the URL between <code>&lt;nowiki&gt;</code> tags as in: <nowiki><nowiki>http://mediawiki.org</nowiki></nowiki> [[Category:Help|Formatting]] 6abff085867e5c83111e48095da28a8290dd2732 1 5 12 2010-01-05T00:26:28Z 192.168.37.29 0 New page: Hello, you can use this page to discuss what's written on the front page. We can also talk on spacefellowship's forum. wikitext text/x-wiki Hello, you can use this page to discuss what's written on the front page. We can also talk on spacefellowship's forum. 490efc5feac3de9e0784a31eb306a4d10f3c6b0f 13 12 2010-01-05T00:27:27Z 192.168.37.29 0 wikitext text/x-wiki Hello, you can use this page to discuss what's written on the front page. We can also talk on [http://spacefellowship.com/Forum/viewforum.php?f=52 Space Fellowship]. d6c6ef2e155c606dd0615efce1cba082db94a893 2 6 14 2010-01-05T01:36:55Z Vincent 1 New page: Greetings, voyager from the stars! I, Vincent, am the administrator of this web site. wikitext text/x-wiki Greetings, voyager from the stars! I, Vincent, am the administrator of this web site. 9330e1782f4d4ed8885b5f8d3f2c5390d33e2b4b 15 14 2010-01-05T01:37:24Z Vincent 1 wikitext text/x-wiki Greetings, voyager from the stars! I, Vincent, am the administrator of this web site. 07cbb92fe3372fc17d47b8861de1a651f8df859b 0 7 17 2010-01-05T02:30:27Z Vincent 1 New page: With the creation of some competitions for private-funded companies/projects in the space domain a few years ago, Dr. Paul H. Dear thought about an other competition open to imaginative pe... wikitext text/x-wiki With the creation of some competitions for private-funded companies/projects in the space domain a few years ago, Dr. Paul H. Dear thought about an other competition open to imaginative people. The idea was [http://www.halfbakery.com/idea/N-Prize brainstormed] and the competition was finally created, along with strict [http://www.n-prize.com rules]. The goal of the competition is to launch from earth surface a very small object into orbit. This satellite must not exceed 20 grams, and the difficult thing is to be able to make this orbital insertion for a cost less than £1,000. There are two sets of rules: the launch can be single use (Single-Spent-to-orbit), like a rocket, and thus using the £1,000 for the full price of the launch vehicle and the satellite. The other way is a reusable launch vehicule (RV) to some extend, for example recovering the first stage of a rocket or a Helium balloon. Everything that is recovered and reusable doesn't count into the £1,000. The pushed-into-orbit [10-20[g satellite has to be tracked by entrants on their own. To be validated, proof must be established that the satellite has made at least 9 orbits before re-entry. 0fce11c6f7fbe661b21a34e78f3c35124cd110b2 12 8 19 2010-01-05T02:44:24Z Vincent 1 Redirecting to [[Help:Formatting]] wikitext text/x-wiki #REDIRECT [[Help:Formatting]] 21281cb3d95e5fe9e6bd0ecae8f68c765f3ec08e 0 9 21 2010-01-22T14:02:31Z Vincent 1 introduction and armadeus wikitext text/x-wiki =Embedded computer: guidance, mission, and telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an intertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positionning, it will require a lot of processing power. An alternative to pure processing power by a CPU is possible: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chipset with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [http://www.armadeus.com/english/products-processor_boards-apf9328.html Armadeus], based on an ARM (FreeScale) processor. ==Software== c27a784642ac05aa7368e490c2c4d83cd9f93cb2 24 21 2010-01-22T18:42:02Z Vincent 1 /* Software */ wikitext text/x-wiki =Embedded computer: guidance, mission, and telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an intertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positionning, it will require a lot of processing power. An alternative to pure processing power by a CPU is possible: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chipset with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [http://www.armadeus.com/english/products-processor_boards-apf9328.html Armadeus], based on an ARM (FreeScale) processor. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [http://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. 1aa5c89d5e98eebe0f32c7d5e827bbb9ec85afa5 33 24 2010-01-25T16:11:00Z Vincent 1 /* Hardware */ xnomai for armadeus wikitext text/x-wiki =Embedded computer: guidance, mission, and telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an intertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positionning, it will require a lot of processing power. An alternative to pure processing power by a CPU is possible: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chipset with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [http://www.armadeus.com/english/products-processor_boards-apf9328.html Armadeus], based on an ARM (FreeScale) processor. Moreover, it [http://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [http://xenomai.org/ Xenomai] Linux-based RTOS. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [http://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. f9df368f70b0fe110223baba3bc8fed548754e6e 39 33 2010-01-26T01:17:53Z Vincent 1 fpga4fun link wikitext text/x-wiki =Embedded computer: guidance, mission, and telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an intertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positionning, it will require a lot of processing power. An alternative to pure processing power by a CPU is possible: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chipset with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [http://www.armadeus.com/english/products-processor_boards-apf9328.html Armadeus], based on an ARM (FreeScale) processor. Moreover, it [http://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [http://xenomai.org/ Xenomai] Linux-based RTOS. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [http://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. 50abf963c4376e8004e2e8a8af5405b3fbc487ac 45 39 2010-01-26T22:05:37Z Vincent 1 /* Hardware */ liens ECB wikitext text/x-wiki =Embedded computer: guidance, mission, and telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an intertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positionning, it will require a lot of processing power. An alternative to pure processing power by a CPU is possible: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chipset with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [http://www.armadeus.com/english/products-processor_boards-apf9328.html Armadeus], based on an ARM (FreeScale) processor. Moreover, it [http://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [http://xenomai.org/ Xenomai] Linux-based RTOS. Other interesting embedded computer boards: the [http://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1] and the [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20]. They do not include a FPGA but are cheaper, the first has a great temperature range, and the second is more powerfull and smaller. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [http://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. 1d934f64656fb27f0b0a40be7e09d9282ab2dddf 6 10 26 2010-01-22T23:44:43Z Vincent 1 Schematic rocket trajectory to explain initial and final direction vectors and atmosphere crossing wikitext text/x-wiki Schematic rocket trajectory to explain initial and final direction vectors and atmosphere crossing b5adc1ae04e61c2c6d9df03f471d2f208f7277ac 28 26 2010-01-22T23:50:39Z Vincent 1 uploaded a new version of "[[Image:Rocket trajectory.png]]" wikitext text/x-wiki Schematic rocket trajectory to explain initial and final direction vectors and atmosphere crossing b5adc1ae04e61c2c6d9df03f471d2f208f7277ac 29 28 2010-01-22T23:52:57Z Vincent 1 uploaded a new version of "[[Image:Rocket trajectory.png]]": earth color change wikitext text/x-wiki Schematic rocket trajectory to explain initial and final direction vectors and atmosphere crossing b5adc1ae04e61c2c6d9df03f471d2f208f7277ac 37 29 2010-01-25T21:26:07Z Vincent 1 uploaded a new version of "[[Image:Rocket trajectory.png]]": lines in white wikitext text/x-wiki Schematic rocket trajectory to explain initial and final direction vectors and atmosphere crossing b5adc1ae04e61c2c6d9df03f471d2f208f7277ac 38 37 2010-01-25T21:28:35Z Vincent 1 uploaded a new version of "[[Image:Rocket trajectory.png]]": croping picture wikitext text/x-wiki Schematic rocket trajectory to explain initial and final direction vectors and atmosphere crossing b5adc1ae04e61c2c6d9df03f471d2f208f7277ac 0 11 41 2010-01-26T19:20:40Z Vincent 1 introducing turbofans wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways is that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. ==General principes== Turbofan blades are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. ==Our design== 09c4d98bea29b34365f741e88b854b2e827b9769 43 41 2010-01-26T19:27:31Z Vincent 1 /* How to build a cheap (~ $150) turbofan? */ wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways is that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. It however has the big disadvantage to be very complicated to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for decision to use turbofans or not. They will have to be build in some way, because it will have to be tailored to our needs, which are quite unusual. ==General principes== Turbofan blades are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. ==Our design== 9592579ccdbb4d6463450214deabb35187a61991 44 43 2010-01-26T19:51:26Z Vincent 1 constraints wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways is that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. It however has the big disadvantage to be very complicated to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for decision to use turbofans or not. They will have to be build in some way, because it will have to be tailored to our needs, which are quite unusual. ==General principes== Turbofan blades are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. ==Our design== Cooling might be needed if low cost metals are used. Startup can be done at ground manually (with compressed air for example). Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. Engine temperature should be used too. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. ed308e1cc6889f1a3a378b04ffa0c81d74c7bdc8 6 12 48 2010-01-29T11:21:29Z Vincent 1 Turbofan principle wikitext text/x-wiki Turbofan principle 91d1cb3d21107df976d7e9aeac0f0d4cdf26a4bb 6 13 49 2010-01-29T11:23:24Z Vincent 1 Turbofan principle 500px wikitext text/x-wiki Turbofan principle 500px 041fe1c80fe7da6216b06fae9cbc20df178573fb 50 49 2010-01-29T11:40:21Z Vincent 1 uploaded a new version of "[[Image:500px-Turbofan operation.svg.png]]": white bg wikitext text/x-wiki Turbofan principle 500px 041fe1c80fe7da6216b06fae9cbc20df178573fb 0 11 51 44 2010-01-29T11:41:06Z Vincent 1 /* General principes */ wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways is that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. It however has the big disadvantage to be very complicated to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for decision to use turbofans or not. They will have to be build in some way, because it will have to be tailored to our needs, which are quite unusual. ==General principes== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan wikipedia's page]. General principe is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion create thrust by evacuating hot gaz. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. It is often seen that a second concentrical shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. [[Image:500px-Turbofan_operation.svg.png]] There are some design properties and configurations that have to be properly calculated depending on the use of the engine, mainly for the intented aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] Turbofan blades are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. ==Our design== Cooling might be needed if low cost metals are used. Startup can be done at ground manually (with compressed air for example). Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. Engine temperature should be used too. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. c2d40ab80d20dc0a09b45bf9a4b215dab7f401a3 52 51 2010-01-29T11:53:50Z Vincent 1 design part adding wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways is that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. It however has the big disadvantage to be very complicated to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for decision to use turbofans or not. They will have to be build in some way, because it will have to be tailored to our needs, which are quite unusual. ==General principes== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan wikipedia's page]. General principe is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion create thrust by evacuating hot gaz. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. It is often seen that a second concentrical shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. [[Image:500px-Turbofan_operation.svg.png]] There are some design properties and configurations that have to be properly calculated depending on the use of the engine, mainly for the intented aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] ==Design versus manufacturing== Design configurations and properties on real engines tend to increase efficiency, meaning higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. ===Shaped core or shaped shaft?=== I think that the main thing that can be optimized to reduce cost and complexity is the design of the compression areas. In the above picture, we see that the shaft is straight and that the core enveloppe is curved to reduce volume on the high compression stage. In practice, the shaft has a bumped profile (small-large-small diameter), to help reduce the volume. ''figure needed'' ===Compressor and turbine blades=== Turbofan blades are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. ==Our design== Cooling might be needed if low cost metals are used. Startup can be done at ground manually (with compressed air for example). Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. Engine temperature should be used too. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. 52211c0e4d7b878c69b8d693aefaeb58b2f5cdb6 55 52 2010-01-29T17:48:00Z Vincent 1 /* Shaped core or shaped shaft? */ image crafted shaft wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways is that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. It however has the big disadvantage to be very complicated to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for decision to use turbofans or not. They will have to be build in some way, because it will have to be tailored to our needs, which are quite unusual. ==General principes== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan wikipedia's page]. General principe is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion create thrust by evacuating hot gaz. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. It is often seen that a second concentrical shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. [[Image:500px-Turbofan_operation.svg.png]] There are some design properties and configurations that have to be properly calculated depending on the use of the engine, mainly for the intented aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] ==Design versus manufacturing== Design configurations and properties on real engines tend to increase efficiency, meaning higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. ===Shaped core or shaped shaft?=== I think that the main thing that can be optimized to reduce cost and complexity is the design of the compression areas. In the above picture, we see that the shaft is straight and that the core enveloppe is curved to reduce volume on the high compression stage. In practice, the shaft has a bumped profile (small-large-small diameter), to help reduce the volume. [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== Turbofan blades are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. ==Our design== Cooling might be needed if low cost metals are used. Startup can be done at ground manually (with compressed air for example). Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. Engine temperature should be used too. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. 3956130d2677e54271bd35ccb69309f8583412a9 56 55 2010-01-29T17:50:52Z Vincent 1 wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways is that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. It however has the big disadvantage to be very complicated to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for decision to use turbofans or not. They will have to be build in some way, because it will have to be tailored to our needs, which are quite unusual. ==General principes== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan wikipedia's page]. General principe is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion create thrust by evacuating hot gaz. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. It is often seen that a second concentrical shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. [[Image:500px-Turbofan_operation.svg.png]] There are some design properties and configurations that have to be properly calculated depending on the use of the engine, mainly for the intented aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] ==Design versus manufacturing== Design configurations and properties on real engines tend to increase efficiency, meaning higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. ===Shaped core or shaped shaft?=== I think that the main thing that can be optimized to reduce cost and complexity is the design of the compression areas. In the above picture, we see that the shaft is straight and that the core enveloppe is curved to reduce volume on the high compression stage. In practice, the shaft has a bumped profile (small-large-small diameter), to help reduce the volume. [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== Turbofan blades are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. Study shape of the blades. Overlapping or not overlapping blades? Single piece of metal with the axis mount ring or assembly of blades on rings? ==Our design== Cooling might be needed if low cost metals are used. Startup can be done at ground manually (with compressed air for example). Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. Engine temperature should be used too. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. efdd13103a158f51969093c2a60a52025f55ec4d 59 56 2010-02-08T01:41:20Z Vincent 1 blades and compression wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways is that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. It however has the big disadvantage to be very complicated to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for decision to use turbofans or not. They will have to be build in some way, because it will have to be tailored to our needs, which are quite unusual. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan wikipedia's page]. General principe is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion create thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] There are some design properties and configurations that have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, meaning higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. ===Shaped core or shaped shaft?=== I think that the main thing that can be optimized to reduce cost and complexity is the design of gas volumes in the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. In practice, the shaft has a bumped profile (small-large-small diameter), to help reduce the volume: [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also on the stator. Without those latter, the air flow would be turning inside the engine, driven by the rotation of the compressor. They allow to redirect the airflow on the next compression stage in the more appropriate and efficient direction. What is the most simple yet efficient shape for turbine, compressor and fan blades? Are a flat shape and a continuous angle acceptable? Overlapping or not overlapping blades? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? For better efficiency, very small gaps should exist between blades and the stator or the rotor for static blades. As always, good efficiency means good precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotation-per-minutes achieved by those engines. ==Our design== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C... Startup can be done at ground manually (with compressed air for example). Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. Engine temperature should be used too. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. 71ad0b6c8d17889dc8b3c2543375735ea9306394 6 14 53 2010-01-29T17:46:14Z Vincent 1 Design proposition for a simpler turbofan. wikitext text/x-wiki Design proposition for a simpler turbofan. 63eb65fcec9f6660fd6b9d94efde46ce77f9afa2 58 53 2010-01-29T19:07:05Z Vincent 1 uploaded a new version of "[[Image:Turbofan craftedshaft.svg]]": updating length and turbine number wikitext text/x-wiki Design proposition for a simpler turbofan. 63eb65fcec9f6660fd6b9d94efde46ce77f9afa2 6 15 54 2010-01-29T17:46:35Z Vincent 1 Design proposition for a simpler turbofan. wikitext text/x-wiki Design proposition for a simpler turbofan. 63eb65fcec9f6660fd6b9d94efde46ce77f9afa2 57 54 2010-01-29T19:06:08Z Vincent 1 uploaded a new version of "[[Image:500px-Turbofan craftedshaft.svg.png]]" wikitext text/x-wiki Design proposition for a simpler turbofan. 63eb65fcec9f6660fd6b9d94efde46ce77f9afa2 0 9 60 45 2010-02-10T00:36:26Z Vincent 1 MCU (ardupilot), telemetry and sensors wikitext text/x-wiki =Embedded computer: guidance, mission, and telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [http://www.armadeus.com/english/products-processor_boards-apf9328.html Armadeus], based on an ARM (FreeScale) processor. Moreover, it [http://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [http://xenomai.org/ Xenomai] Linux-based RTOS. Other interesting embedded computer boards: the [http://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1] and the [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20]. They do not include a FPGA but are cheaper, the first has a great temperature range, and the second is more powerful and smaller. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [http://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. ===Telemetry=== It seems that the 900MHz version of the ZigBee communication standard is able to transmit at around 100kbps up to 10km. Taken from th ArduPilot page: Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [http://www.adafruit.com/index.php?main_page=product_info&cPath=29&products_id=126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [http://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [http://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * accelerometer * [http://store.diydrones.com/ProductDetails.asp?ProductCode=SE-0002-01 XYZ horizon sensor] * thermometer * altimeter / is it actually possible? probably with accelerometer * 3D compass * fuel gauge or low level indicator ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [http://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. d6acc969555aded460073ea1e6b165eb16a74566 78 60 2010-10-17T19:59:05Z Vincent 1 sensors wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [http://www.armadeus.com/english/products-processor_boards-apf9328.html Armadeus], based on an ARM (FreeScale) processor. Moreover, it [http://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [http://xenomai.org/ Xenomai] Linux-based RTOS. Other interesting embedded computer boards: the [http://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1] and the [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20]. They do not include a FPGA but are cheaper, the first has a great temperature range, and the second is more powerful and smaller. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [http://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. ===Telemetry=== It seems that the 900MHz version of the ZigBee communication standard is able to transmit at around 100kbps up to 10km. Taken from th ArduPilot page: Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [http://www.adafruit.com/index.php?main_page=product_info&cPath=29&products_id=126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [http://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [http://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: quite common nowadays, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the theoretically computed strength the mechanical structure has to sustain. To chose a sensor, sparkfun wrote an [http://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [http://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and it might be replaced by the camera sensor we want to develop. * Magnetometer (3D compass): lots of sensors exist too, for example the |http://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * [http://store.diydrones.com/ProductDetails.asp?ProductCode=SE-0002-01 XYZ Horizon sensor] * Thermometer: for systems health monitoring, like engines temperature. * GPS? if USAF allows it in altitude... * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [http://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. 9d1c4dda541310226f04bfd518802876fcbb58ce 79 78 2010-10-23T09:56:11Z Vincent 1 Failsafe wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [http://www.armadeus.com/english/products-processor_boards-apf9328.html Armadeus], based on an ARM (FreeScale) processor. Moreover, it [http://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [http://xenomai.org/ Xenomai] Linux-based RTOS. Other interesting embedded computer boards: the [http://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1] and the [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20]. They do not include a FPGA but are cheaper, the first has a great temperature range, and the second is more powerful and smaller. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [http://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. ===Telemetry=== It seems that the 900MHz version of the ZigBee communication standard is able to transmit at around 100kbps up to 10km. Taken from th ArduPilot page: Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [http://www.adafruit.com/index.php?main_page=product_info&cPath=29&products_id=126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [http://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [http://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: quite common nowadays, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the theoretically computed strength the mechanical structure has to sustain. To chose a sensor, sparkfun wrote an [http://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [http://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and it might be replaced by the camera sensor we want to develop. * Magnetometer (3D compass): lots of sensors exist too, for example the |http://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * [http://store.diydrones.com/ProductDetails.asp?ProductCode=SE-0002-01 XYZ Horizon sensor] * Thermometer: for systems health monitoring, like engines temperature. * GPS? if USAF allows it in altitude... * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [http://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. 93e77e3b1d7b82df948c573b4d9474ff83d8a6b4 85 79 2010-10-23T14:07:25Z Vincent 1 /* High processing power */ armadeus page link. wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [http://www.armadeus.com/english/products-processor_boards-apf9328.html Armadeus], based on an ARM (FreeScale) processor. Moreover, it [http://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [http://xenomai.org/ Xenomai] Linux-based RTOS. Armadeus board integration has a [[Armadeus|dedicated page]]. Other interesting embedded computer boards: the [http://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1] and the [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20]. They do not include a FPGA but are cheaper, the first has a great temperature range, and the second is more powerful and smaller. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [http://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. ===Telemetry=== It seems that the 900MHz version of the ZigBee communication standard is able to transmit at around 100kbps up to 10km. Taken from th ArduPilot page: Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [http://www.adafruit.com/index.php?main_page=product_info&cPath=29&products_id=126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [http://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [http://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: quite common nowadays, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the theoretically computed strength the mechanical structure has to sustain. To chose a sensor, sparkfun wrote an [http://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [http://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and it might be replaced by the camera sensor we want to develop. * Magnetometer (3D compass): lots of sensors exist too, for example the |http://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * [http://store.diydrones.com/ProductDetails.asp?ProductCode=SE-0002-01 XYZ Horizon sensor] * Thermometer: for systems health monitoring, like engines temperature. * GPS? if USAF allows it in altitude... * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [http://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. 8804d69997cd080be5eab516bb0640a3a0cbf82d 0 1 61 42 2010-04-27T01:20:30Z Vincent 1 JPaerospace + SSTO wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one], here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but crusing at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerozene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, and earth curve recognition from 30km altitude can be used for that too. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has tons of drawbacks because of the need for cryogenics storage, manipulation, and engine design, that make it quite expensive and too much complicated. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... f2c4f2416c5be060c8429f6340e9f214177ca6ec 62 61 2010-04-27T01:23:37Z Vincent 1 /* Guidance */ camera wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one], here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but crusing at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerozene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has tons of drawbacks because of the need for cryogenics storage, manipulation, and engine design, that make it quite expensive and too much complicated. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... a8b1f91cb2ed50062439f865cc650d091042cf81 63 62 2010-06-18T00:31:16Z Vincent 1 piston pump link wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one], here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but crusing at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerozene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has tons of drawbacks because of the need for cryogenics storage, manipulation, and engine design, that make it quite expensive and too much complicated. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is used on a 1,500 lb-thrust LOX/kerosene engine. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 15efeee6201e1cca58c9bef4ba021cb16c50b4ba 64 63 2010-06-18T00:32:36Z Vincent 1 cryogenic engineering book wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally roughly form a square angle, with the beginning of the flight is orthogonal to Earth and the final direction is parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) is quite important. It is more efficient to first escape the atmosphere and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbo design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasability of something similar, at low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one], here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but crusing at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from high speeds of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerozene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's onboard computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost. See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Ergols represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the german V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerozene (RP-1) has been introduced later to replace alcohol, providing a beter volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has tons of drawbacks because of the need for cryogenics storage, manipulation, and engine design, that make it quite expensive and too much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. Webpage on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I only see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as an option. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is used on a 1,500 lb-thrust LOX/kerosene engine. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket. Both case imply lot of complications of the rocket's hardware. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, narrowband EM emitting, mirrors... 4b19de5e7e5773a61e29f265c59bda310a5c74c8 70 64 2010-07-26T21:15:15Z Vincent 1 typos wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines should be seriously considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as the only option. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. eadde14a316bc9f1f5d4e9139bf3e418d83d1141 71 70 2010-08-03T11:56:05Z Vincent 1 /* Engine */ wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that E85, a 85 percent alcohol and 15 percent gasoline fuel recently used in automotive, is promising. I think that it's efficiency will be slightly better than alcohol, still being very cheap, around £0.5 a liter. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as the only option. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 65f8cdacb12a4628dcd413bfc517bd667a276030 8 2 65 47 2010-06-27T22:56:37Z Vincent 1 adding resources wikitext text/x-wiki * navigation ** mainpage|mainpage ** Resources|Resources ** recentchanges-url|recentchanges ** randompage-url|randompage * sections ** EmbeddedRocketComputer|Computer ** Build_a_cheap_turbofan|Plane engine 3fd400cb118434454dce644de5f2364fd4d57146 83 65 2010-10-23T14:04:01Z Vincent 1 adding testing section wikitext text/x-wiki * navigation ** mainpage|mainpage ** Resources|Resources ** recentchanges-url|recentchanges ** randompage-url|randompage * sections ** EmbeddedRocketComputer|Computer ** Build_a_cheap_turbofan|Plane engine ** Testing and validation|Testing 7f1265b1666c2f168b59c0f19f530d5339acb4a3 84 83 2010-10-23T14:04:27Z Vincent 1 wikitext text/x-wiki * navigation ** mainpage|mainpage ** Resources|Resources ** recentchanges-url|recentchanges ** randompage-url|randompage * sections ** EmbeddedRocketComputer|Computer ** Build_a_cheap_turbofan|Plane engine ** Testing|Testing and validation 789b594cb4e88e56741695637b98f5138cad076d 0 16 66 2010-06-27T23:20:16Z Vincent 1 Adding resources pages: google book links on turbines and rocket engines wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Books (online links)== ===Multi-domain=== * [http://books.google.fr/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://books.google.fr/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.fr/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. ====Specific parts:==== * [http://books.google.fr/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.fr/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.fr/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.fr/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.fr/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.fr/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos== eba6ed0ab900fca7400bd048774ffe22bc2ad97e 67 66 2010-06-27T23:24:42Z Vincent 1 /* Videos */ adding youtube links for ssto, how it's made, and jpa wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Books (online links)== ===Multi-domain=== * [http://books.google.fr/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://books.google.fr/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.fr/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. ====Specific parts:==== * [http://books.google.fr/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.fr/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.fr/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.fr/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.fr/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.fr/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. db41b00e96307b78364df7d88cd3ea20544c5daa 68 67 2010-06-27T23:33:04Z Vincent 1 /* Resources */ adding web pages (xcor only) wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Web pages== * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. ==Books (online links)== ===Multi-domain=== * [http://books.google.fr/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://books.google.fr/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.fr/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. ====Specific parts:==== * [http://books.google.fr/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.fr/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.fr/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.fr/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.fr/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.fr/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. 603e99e45c9971e675fbbd0156aba08e4038d061 69 68 2010-06-27T23:38:50Z Vincent 1 Adding link on nozzle design and cryo book wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Web pages== * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. ==Books (online links)== ===Multi-domain=== * [http://books.google.fr/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://books.google.fr/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.fr/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. ====Specific parts:==== * [http://books.google.fr/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.fr/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.fr/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.fr/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.fr/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.fr/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.fr/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. 81ef7066aef3f7ef3bd89b6cc7e82c21fe3a10ae 98 69 2010-11-17T08:39:54Z Vincent 1 /* Web pages */ add r7 link wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Web pages== * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm history of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Books (online links)== ===Multi-domain=== * [http://books.google.fr/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://books.google.fr/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.fr/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. ====Specific parts:==== * [http://books.google.fr/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.fr/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.fr/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.fr/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.fr/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.fr/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.fr/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. d9f081dd5a4240c805a2341ccbf08ee79e0c9841 99 98 2010-11-17T08:40:11Z Vincent 1 /* Web pages */ wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Web pages== * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Books (online links)== ===Multi-domain=== * [http://books.google.fr/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://books.google.fr/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.fr/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. ====Specific parts:==== * [http://books.google.fr/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.fr/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.fr/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.fr/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.fr/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.fr/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.fr/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. 3f26957941da868357b3e71ae85b55eae170d163 0 17 72 2010-08-03T12:09:56Z Vincent 1 Table begin wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''FRocket engines features'' |- |'''Company''' !NASA !XCOR !Virgin |- |'''Model''' |SSME | | |- |'''Propellers''' |LOX & LH2 |LOX & alcohol |LOX & kerosene |- |'''Tank pressurization''' |Yes | - |Yes |- |'''Fuel pump''' |Turbopump |Piston pump |? |- |'''Cooling''' |Regenerative |Regenerative |Regenerative |- |'''Chamber metal''' |? |Copper |? |- |} 855f70e4f1a0c201b8a898f2198398a30433d653 73 72 2010-08-03T14:01:18Z Vincent 1 /* Rocket Engine */ wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''FRocket engines features'' |- |'''Company''' !Rocketdyne !XCOR !XCOR !Virgin |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |SpaceShipOne |- |'''Propellers''' |LOX & LH2 |LOX & Alcohol |LOX & Kerosene |LOX & Kerosene |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |No |No |Yes |- |'''Fuel pump''' |Turbopump |Piston pump |Piston pump |? |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative |Regenerative w/ Kerosene |Regenerative |- |'''Chamber metal''' |Copper or iron? |Copper |Copper |? |- |} fc8bb99b6a0238d77891e40921d966f75905784b 74 73 2010-08-03T15:09:55Z Vincent 1 wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !XCOR !XCOR |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |- |'''Propellers''' |LOX & LH2 |LOX & Alcohol |LOX & Kerosene |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |No |No |- |'''Fuel pump''' |Turbopump |Piston pump |Piston pump |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |- |'''Chamber metal''' |Copper or iron? |Copper |Copper |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Traditionnaly, only turbo pumps have been able to feed the engine at a large enough rate. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. ==Cooling== Regenerative cooling is most widely used in rocket engines. Few of them however use other ways, like ablatively cooling carbon fiber composite in SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine, or radiative cooling in the Merlin Vacuum nozzle (still regenerative for the chamber). ea8473ab15e725f3056482872485713d0ba92884 75 74 2010-08-03T15:16:02Z Vincent 1 /* Rocket Engine */ wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !XCOR !XCOR |Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- |'''Propellers''' |LOX & LH2 |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |No |No |Yes, with Helium |- |'''Fuel pump''' |Turbopump |Piston pump |Piston pump |No |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |'''Chamber metal''' |Copper or iron? |Copper |Copper |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Traditionnaly, only turbo pumps have been able to feed the engine at a large enough rate. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. ==Cooling== Regenerative cooling is most widely used in rocket engines. Few of them however use other ways, like ablatively cooling carbon fiber composite in SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine, or radiative cooling in the Merlin Vacuum nozzle (still regenerative for the chamber). 363f7e12ff643418e1903305ea5a9a3efe08ec0f 76 75 2010-09-20T22:16:08Z Vincent 1 /* Rocket Engine */ more rows wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellers''' |LOX & LH2 |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |No |No |Yes, with Helium |- |'''Fuel pump''' |Turbopump |Piston pump |Piston pump |No |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |'''Chamber metal''' |Copper or iron? |Copper |Copper |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | |Electric |- |'''Provided by''' |Engine's turbopumps | | |? |- |'''Actuator''' |six hydraulic servoactuators |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically actuated |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Traditionnaly, only turbo pumps have been able to feed the engine at a large enough rate. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. ==Cooling== Regenerative cooling is most widely used in rocket engines. Few of them however use other ways, like ablatively cooling carbon fiber composite in SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine, or radiative cooling in the Merlin Vacuum nozzle (still regenerative for the chamber). fffe51955f14ded218e0946ef9891d2f964a7f93 77 76 2010-09-20T23:53:12Z Vincent 1 /* Rocket Engine */ wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellers''' |LOX & LH2 |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |No |No |Yes, with Helium |- |'''Fuel pump''' |Turbopump |Piston pump |Piston pump |No |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |'''Chamber metal''' |Copper or iron? |Copper |Copper |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | |Electric |- |'''Provided by''' |Engine's turbopumps | | |? |- |'''Actuator''' |six hydraulic servoactuators |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Traditionnaly, only turbo pumps have been able to feed the engine at a large enough rate. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. ==Cooling== Regenerative cooling is most widely used in rocket engines. Few of them however use other ways, like ablatively cooling carbon fiber composite in SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine, or radiative cooling in the Merlin Vacuum nozzle (still regenerative for the chamber). d31328cee59af8fe33d1bd753762b74e71665073 95 77 2010-11-12T22:30:13Z Vincent 1 typo fix wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |No |No |Yes, with Helium |- |'''Fuel pump''' |Turbopump |Piston pump |Piston pump |No |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |'''Chamber metal''' |Copper or iron? |Copper |Copper |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | |Electric |- |'''Provided by''' |Engine's turbopumps | | |? |- |'''Actuator''' |six hydraulic servoactuators |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Traditionnaly, only turbo pumps have been able to feed the engine at a large enough rate. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. ==Cooling== Regenerative cooling is most widely used in rocket engines. Few of them however use other ways, like ablatively cooling carbon fiber composite in SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine, or radiative cooling in the Merlin Vacuum nozzle (still regenerative for the chamber). 81c3cd458299e90881f58b847821aab2e7818cb4 96 95 2010-11-17T08:24:50Z Vincent 1 New colon on soyuz engine and new ignition row. wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series (Soyuz) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |Yes, with Nitrogen |No |No |Yes, with Helium |- |'''Fuel pump''' |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition |Piston pump |Piston pump |No |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative? |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |'''Chamber metal''' |Copper or iron? |Copper? |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic |? | | |Electric |- |'''Provided by''' |Engine's turbopumps |? | | |? |- |'''Actuator''' |six hydraulic servoactuators |? |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Traditionnaly, only turbo pumps have been able to feed the engine at a large enough rate. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. ==Cooling== Regenerative cooling is most widely used in rocket engines. Few of them however use other ways, like ablatively cooling carbon fiber composite in SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine, or radiative cooling in the Merlin Vacuum nozzle (still regenerative for the chamber). 9dbe09e0a5d73e2b7ce321481ebac3d137c97575 97 96 2010-11-17T08:38:40Z Vincent 1 /* Rocket Engine */ cooling and more info on soyuz wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |Yes, with Nitrogen |No |No |Yes, with Helium |- |'''Fuel pump''' |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |'''Injector''' |? |337 swirling/mixing injectors, ring of kerosene only for film cooling |? |? |? |- |'''Chamber metal''' |Copper or iron? |Copper sheet inside, steel outside? |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic |? | | |Electric |- |'''Provided by''' |Engine's turbopumps |? | | |? |- |'''Actuator''' |Six hydraulic servoactuators |? |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Traditionnaly, only turbo pumps have been able to feed the engine at a large enough rate. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. ==Cooling== Regenerative cooling is most widely used in rocket engines. Few of them however use other ways, like ablatively cooling carbon fiber composite in SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine, or radiative cooling in the Merlin Vacuum nozzle (still regenerative for the chamber). d0a7e6f5a6fdc41ac70e47de38b39e84c876e54c 0 18 80 2010-10-23T10:18:35Z Vincent 1 Beginning of testing wikitext text/x-wiki =Testing= Once created, systems and subsystems will have to be carefully tested, checking for defects, design errors, model predictions, programming errors, and performance. In this page, we detail what will have to be tested for each subsystem, and how it can be done with a minimal cost. ==Structure== ===Wings=== Wings will have to sustain the mass of the rocket at accelerations from -2g to +4g. They will have to support heavy curving, and this can be tested by fixing the joint and putting a weight on the edge of wings. This weight will be calculated from the lift, mass and maximal acceleration that has to be endured. ===Rocket body=== Rocket body constraints are quite hard to calculate. Air pressure is the main problem, and happens in two ways: * on the fairing during normal operation due to speed, * on the body, when actuating the engine (or other) to modify pitch or yaw of the rocket. This force can be pretty intense when speed is mach3. Fortunately, launching from a plane will reduce air density drastically. ===Joint between wings and body=== The wings and the body have to be tightly coupled, to sustain the inertia of the heavy body. It also has to be simple enough to be separated for staging. ==Mechanical systems== ===Plane actuators (for ailerons, elevators...)=== ===Fuel pumps=== ===Staging mechanism=== ==Plane engines== ===Fans=== ===Physical property=== Thrust, temperature. Long-term running validation (150% of available fuel at least). ==Telemetry== ==Rocket engine== cf1dcd4a2bf7766ac5ecc9a073d2d9d4524d1eee 81 80 2010-10-23T11:00:56Z Vincent 1 more testing wikitext text/x-wiki =Testing= Once created, systems and subsystems will have to be carefully tested, checking for defects, design errors, model predictions, programming errors, and performance. In this page, we detail what will have to be tested for each subsystem, and how it can be done with a minimal cost. ==Structure== ===Wings=== Wings will have to sustain the mass of the rocket at accelerations from -2g to +4g. They will have to support heavy curving, and this can be tested by fixing the joint and putting a weight on the edge of wings. This weight will be calculated from the lift, mass and maximal acceleration that has to be endured. ===Rocket body=== Rocket body constraints are quite hard to calculate. Air pressure is the main problem, and happens in two ways: * on the fairing during normal operation due to speed, * on the body, when actuating the engine (or other) to modify pitch or yaw of the rocket. This force can be pretty intense when speed is mach3. Fortunately, launching from a plane will reduce air density drastically. ===Joint between wings and body=== The wings and the body have to be tightly coupled, to sustain the inertia of the heavy body. It also has to be simple enough to be separated for staging. ==Mechanical systems== ===Plane actuators (for ailerons, elevators...)=== We'll have to make sure the selected actuators conform to their specification, for the torque at least. ===Fuel pumps=== Fuel pumps are a critical part of the thrust system. If they don't provide a constant throughput, it's very likely that something will fail, especially for the rocket engine. Flow must not vary with accelerations of the liquid inside the tanks for example. If a pressurization system is preferred over pumps for the rocket propellants, we'll have to make sure that there are no interruptions in the flow either, due to accelerations withstood. ===Staging mechanism=== The mechanism will have to be tested in full load operations, with strong accelerations (more weight). The staging sensor is a key element, because it will trigger the mission continuation, so we need to fully qualify it to know if staging could have occurred even if it did not detect it, or the opposite. ==Plane engines== ===Fans=== Fans, and the fan shaft, will need to be carefully checked for balancing. This can be simply done with vibration sensing at various rotation speeds. Proper frequency of all the fans will thus be evaluated, allowing to model the behavior of all moving parts of the engine. It however does not allow to fully model the engine, because the frequencies may change because of the air displacement and pressure effects caused by the (core) shell of the engine. The final assembly can be checked in the same way than the blade, by checking vibrations at various rotation speeds. A key element in turbine engines is the loose between the blades and the stator. They have to be really close one of each other for the engine performs efficiently. It means that the blades have to be exactly the same size (even if the manufacturing process of each blade is based on that, the assembly of a fan should be checked), and an abrasive crafting of the stator can be done. Revolving the rotor inside a soft material stator is a very simple and efficient process to ensure a minimal loose. ===Physical property=== Thrust, temperature. Long-term running validation (150% of available fuel at least). ==Telemetry== ==Rocket engine== 8ac507b4a2b0b8a71510b2d6cbf7d325ba40773b 82 81 2010-10-23T11:52:38Z Vincent 1 more testing wikitext text/x-wiki =Testing= Once created, systems and subsystems will have to be carefully tested, checking for defects, design errors, model predictions, programming errors, and performance. In this page, we detail what will have to be tested for each subsystem, and how it can be done with a minimal cost. ==Structure== ===Wings=== Wings will have to sustain the mass of the rocket at accelerations from -2g to +4g. They will have to support heavy curving, and this can be tested by fixing the joint and putting a weight on the edge of wings. This weight will be calculated from the lift, mass and maximal acceleration that has to be endured. ===Rocket body=== Rocket body constraints are quite hard to calculate. Air pressure is the main problem, and happens in two ways: * on the fairing during normal operation due to speed, * on the body, when actuating the engine (or other) to modify pitch or yaw of the rocket. This force can be pretty intense when speed is mach3. Fortunately, launching from a plane will reduce air density drastically. ===Joint between wings and body=== The wings and the body have to be tightly coupled, to sustain the inertia of the heavy body. It also has to be simple enough to be separated for staging. ==Mechanical systems== ===Plane actuators (for ailerons, elevators...)=== We'll have to make sure the selected actuators conform to their specification, for the torque at least. ===Fuel pumps=== Fuel pumps are a critical part of the thrust system. If they don't provide a constant throughput, it's very likely that something will fail, especially for the rocket engine. Flow must not vary with accelerations of the liquid inside the tanks for example. If a pressurization system is preferred over pumps for the rocket propellants, we'll have to make sure that there are no interruptions in the flow either, due to accelerations withstood. ===Staging mechanism=== The mechanism will have to be tested in full load operations, with strong accelerations (more weight). The staging sensor is a key element, because it will trigger the mission continuation, so we need to fully qualify it to know if staging could have occurred even if it did not detect it, or the opposite. ===Fairing=== A system with pre-tensioned springs should be the simplest option. ===Orbital insertion=== A simple orbital insertion mechanism is the releases of pre-tensioned springs, like what's used in Ariane rockets. However, they release the springs with a pyrotechnic system, and we'll have to find something less dangerous, costly and more easy to build. ==Plane engines== ===Fans=== Fans, and the fan shaft, will need to be carefully checked for balancing. This can be simply done with vibration sensing at various rotation speeds. Proper frequency of all the fans will thus be evaluated, allowing to model the behavior of all moving parts of the engine. It however does not allow to fully model the engine, because the frequencies may change because of the air displacement and pressure effects caused by the (core) shell of the engine. The final assembly can be checked in the same way than the blade, by checking vibrations at various rotation speeds. A key element in turbine engines is the loose between the blades and the stator. They have to be really close one of each other for the engine performs efficiently. It means that the blades have to be exactly the same size (even if the manufacturing process of each blade is based on that, the assembly of a fan should be checked), and an abrasive crafting of the stator can be done. Revolving the rotor inside a soft material stator is a very simple and efficient process to ensure a minimal loose. ===Physical property=== Thrust, temperature. Long-term running validation (150% of available fuel at least). ==Telemetry== The telemetry will only be able to work to these distances because there are no obstacles but the atmosphere. Therefore, testing it on real distance on the floor is mission impossible. The best way may be to measure up to how much attenuation the telemetry still works, and check if this is the expected attenuation of the atmosphere. For tracking, tests should be made with a simple RC plane embedding our emitter. This will allow us to model the cone (or other shape) of reception with regards to optimal gain settings. It will also allow to check the ability to point the directive antenna to a (too quickly) moving object. ==Rocket engine== TBD ==Pre-prototype sensor validation== Some tests will have to be run before actually designing the whole system, in order to validate the use of a material, the capabilities of a particular part, or a design choice. Most sensor validation will be done using a simple RC plane. ===GPS=== GPS tracking seems to be not accurate and not suitable for altitude sensing. This may be a limitation of the chips, because some satellites use GPS to position themselves, so it is not a system limitation. ===Magnetometer=== Embedding a 3-axis magnetometer on the plane will allow us to check if this information can be used and with what accuracy. 125685d07fbaf9916eac3729f014d34e4b26d9c5 0 19 86 2010-10-23T14:15:01Z Vincent 1 Beginning of Armadeus page: RTC wikitext text/x-wiki =Armadeus single circuit board= ==Integration board== The Armadeus board has two big connectors (Hirose type) and need to be plugged on a main board through these two connectors to communicate with the outside world, and more simply, with electrical power. Armadeus developed several development and test boards. They include lots of input/output capabilities, power management, and so on. For our goal, we need to develop a new main board, with a reduced size and weight compared to the Armadeus' development board, and with only the needed external connectors and components. On this page, we make the list of those needed connectors and design constraints. ===Real-time clock=== The RTC device is used to keep track of the time in case of power interruption or reboot. The DevFull board from Armadeus embeds one, the [http://www.maxim-ic.com/datasheet/index.mvp/id/3806 Maxim DS1374]. It uses the I²C bus and has an internal temperature-compensated oscillator. Datasheet is uploaded on this page. a0511428f18fd8108b796c76024f2a765f75a397 89 86 2010-10-23T14:24:41Z Vincent 1 /* Real-time clock */ link to datasheet wikitext text/x-wiki =Armadeus single circuit board= ==Integration board== The Armadeus board has two big connectors (Hirose type) and need to be plugged on a main board through these two connectors to communicate with the outside world, and more simply, with electrical power. Armadeus developed several development and test boards. They include lots of input/output capabilities, power management, and so on. For our goal, we need to develop a new main board, with a reduced size and weight compared to the Armadeus' development board, and with only the needed external connectors and components. On this page, we make the list of those needed connectors and design constraints. ===Real-time clock=== The RTC device is used to keep track of the time in case of power interruption or reboot. The DevFull board from Armadeus embeds one, the [http://www.maxim-ic.com/datasheet/index.mvp/id/3806 Maxim DS1374]. It uses the I²C bus and has an internal temperature-compensated oscillator. Datasheet is uploaded [[media:Datasheet_DS1374.pdf|here]]. f935aa447e14be21283df463302b0977bd45d9ae 90 89 2010-10-23T14:38:21Z Vincent 1 More devices to be integrated. wikitext text/x-wiki =Armadeus single circuit board= ==Integration board== The Armadeus board has two big connectors (Hirose 120 pin [http://www.hirose-connectors.com/connectors/H205SeriesGaiyou.aspx?c1=FX8&c3=3 FX8]) and need to be plugged on a main board through these two connectors to communicate with the outside world, and more simply, with electrical power. Armadeus developed a few [http://www.armadeus.com/english/products-development_boards-apf27_dev.html development boards]. They include lots of input/output capabilities, power management, and so on. For our goal, we need to develop a new main board, with a reduced size and weight compared to the Armadeus' development board, and with only the needed external connectors and components. On this page, we make the list of those needed connectors and design constraints. ===Real-time clock=== The RTC device is used to keep track of the time in case of power interruption or reboot. The DevFull board from Armadeus embeds one, the [http://www.maxim-ic.com/datasheet/index.mvp/id/3806 Maxim DS1374]. It uses the I²C bus and has an internal temperature-compensated oscillator. Datasheet is uploaded [[media:Datasheet_DS1374.pdf|here]]. ===SD card=== We might need an SD card for massive storage of sensor and flight information. The default is 256MB on the [http://www.armadeus.com/english/products-processor_boards-apf27.html APF27] processor board. ===ADC=== An ADC may be needed if analogic sensors are used. On the Armadeus DevFull board, they use the MAX1027 component. ===Camera module=== Camera modules can't be directly connected to APF signals, but need a breakout board. Moreover, since they need to be put in a different place than the processing board, they need a connector on the main board. 42a02f540c124554853984425ebd56b771a8df88 91 90 2010-10-23T14:44:40Z Vincent 1 /* Integration board */ PSU voltages wikitext text/x-wiki =Armadeus single circuit board= ==Integration board== The Armadeus board has two big connectors (Hirose 120 pin [http://www.hirose-connectors.com/connectors/H205SeriesGaiyou.aspx?c1=FX8&c3=3 FX8]) and need to be plugged on a main board through these two connectors to communicate with the outside world, and more simply, with electrical power. Armadeus developed a few [http://www.armadeus.com/english/products-development_boards-apf27_dev.html development boards]. They include lots of input/output capabilities, power management, and so on. For our goal, we need to develop a new main board, with a reduced size and weight compared to the Armadeus' development board, and with only the needed external connectors and components. On this page, we make the list of those needed connectors and design constraints. ===Power supply=== The dev board embeds three converters and regulators, for 1.8V, 3.3V and 5V. ===Real-time clock=== The RTC device is used to keep track of the time in case of power interruption or reboot. The DevFull board from Armadeus embeds one, the [http://www.maxim-ic.com/datasheet/index.mvp/id/3806 Maxim DS1374]. It uses the I²C bus and has an internal temperature-compensated oscillator. Datasheet is uploaded [[media:Datasheet_DS1374.pdf|here]]. ===SD card=== We might need an SD card for massive storage of sensor and flight information. The default is 256MB on the [http://www.armadeus.com/english/products-processor_boards-apf27.html APF27] processor board. ===ADC=== An ADC may be needed if analogic sensors are used. On the Armadeus DevFull board, they use the MAX1027 component. ===Camera module=== Camera modules can't be directly connected to APF signals, but need a breakout board. Moreover, since they need to be put in a different place than the processing board, they need a connector on the main board. b5333cdb8b6bd40b96dd87b92dd6a7436e7942f5 92 91 2010-10-23T14:48:16Z Vincent 1 /* Armadeus single circuit board */ wikitext text/x-wiki =Armadeus small processor boards= ==Integration board== The Armadeus board has two big connectors (Hirose 120 pin [http://www.hirose-connectors.com/connectors/H205SeriesGaiyou.aspx?c1=FX8&c3=3 FX8]) and need to be plugged on a main board through these two connectors to communicate with the outside world, and more simply, with electrical power. Armadeus developed a few [http://www.armadeus.com/english/products-development_boards-apf27_dev.html development boards]. They include lots of input/output capabilities, power management, and so on. For our goal, we need to develop a new main board, with a reduced size and weight compared to the Armadeus' development board, and with only the needed external connectors and components. On this page, we make the list of those needed connectors and design constraints. ===Power supply=== The dev board embeds three converters and regulators, for 1.8V, 3.3V and 5V. ===Real-time clock=== The RTC device is used to keep track of the time in case of power interruption or reboot. The DevFull board from Armadeus embeds one, the [http://www.maxim-ic.com/datasheet/index.mvp/id/3806 Maxim DS1374]. It uses the I²C bus and has an internal temperature-compensated oscillator. Datasheet is uploaded [[media:Datasheet_DS1374.pdf|here]]. ===SD card=== We might need an SD card for massive storage of sensor and flight information. The default is 256MB on the [http://www.armadeus.com/english/products-processor_boards-apf27.html APF27] processor board. ===ADC=== An ADC may be needed if analogic sensors are used. On the Armadeus DevFull board, they use the MAX1027 component. ===Camera module=== Camera modules can't be directly connected to APF signals, but need a breakout board. Moreover, since they need to be put in a different place than the processing board, they need a connector on the main board. 42bf31ed5865c8b33f48b81b2bb1c78b6f6b155b 93 92 2010-10-23T15:00:22Z Vincent 1 FPGA for PWM wikitext text/x-wiki =Armadeus small processor boards= ==Integration board== The Armadeus board has two big connectors (Hirose 120 pin [http://www.hirose-connectors.com/connectors/H205SeriesGaiyou.aspx?c1=FX8&c3=3 FX8]) and need to be plugged on a main board through these two connectors to communicate with the outside world, and more simply, with electrical power. Armadeus developed a few [http://www.armadeus.com/english/products-development_boards-apf27_dev.html development boards]. They include lots of input/output capabilities, power management, and so on. For our goal, we need to develop a new main board, with a reduced size and weight compared to the Armadeus' development board, and with only the needed external connectors and components. On this page, we make the list of those needed connectors and design constraints. ===Power supply=== The dev board embeds three converters and regulators, for 1.8V, 3.3V and 5V. ===Real-time clock=== The RTC device is used to keep track of the time in case of power interruption or reboot. The DevFull board from Armadeus embeds one, the [http://www.maxim-ic.com/datasheet/index.mvp/id/3806 Maxim DS1374]. It uses the I²C bus and has an internal temperature-compensated oscillator. Datasheet is uploaded [[media:Datasheet_DS1374.pdf|here]]. ===SD card=== We might need an SD card for massive storage of sensor and flight information. The default is 256MB on the [http://www.armadeus.com/english/products-processor_boards-apf27.html APF27] processor board. ===ADC=== An ADC may be needed if analogic sensors are used. On the Armadeus DevFull board, they use the MAX1027 component. ===Camera module=== Camera modules can't be directly connected to APF signals, but need a breakout board. Moreover, since they need to be put in a different place than the processing board, they need a connector on the main board. ===FPGA connectors for servo PWM=== If we chose servo as the actuators for the airplane ailerons, we will need a PWM signal to control them. The FPGA of Armadeus boards is able to control lots of them as explained [http://www.fpga4fun.com/RCServos.html here], but we need to add the connectors to the main board. 1de7ad5c88b04efc1c1b566349ccac18a1574715 94 93 2010-10-23T15:12:59Z Vincent 1 /* Real-time clock */ wikitext text/x-wiki =Armadeus small processor boards= ==Integration board== The Armadeus board has two big connectors (Hirose 120 pin [http://www.hirose-connectors.com/connectors/H205SeriesGaiyou.aspx?c1=FX8&c3=3 FX8]) and need to be plugged on a main board through these two connectors to communicate with the outside world, and more simply, with electrical power. Armadeus developed a few [http://www.armadeus.com/english/products-development_boards-apf27_dev.html development boards]. They include lots of input/output capabilities, power management, and so on. For our goal, we need to develop a new main board, with a reduced size and weight compared to the Armadeus' development board, and with only the needed external connectors and components. On this page, we make the list of those needed connectors and design constraints. ===Power supply=== The dev board embeds three converters and regulators, for 1.8V, 3.3V and 5V. ===Real-time clock=== The RTC device is used to keep track of the time in case of power interruption or reboot. The processor board embeds one, but it is not saved in case of power interruption. Consequently, the DevFull board from Armadeus embeds another one, the [http://www.maxim-ic.com/datasheet/index.mvp/id/3806 Maxim DS1374] with a battery. It uses the I²C bus and has an internal temperature-compensated oscillator. Datasheet is uploaded [[media:Datasheet_DS1374.pdf|here]]. For our main board, we only need the battery, since the main board and the processor board should not be decoupled. Using a GPS would be a benefit since it provides accurate time. Else, upload telemetry has to be setup to allow time setting from the outside of the system, when the board is plugged to the main board. If this is too boring, putting a RTC on the main board, as Armadeus does, can be considered. ===SD card=== We might need an SD card for massive storage of sensor and flight information. The default is 256MB on the [http://www.armadeus.com/english/products-processor_boards-apf27.html APF27] processor board. ===ADC=== An ADC may be needed if analogic sensors are used. On the Armadeus DevFull board, they use the MAX1027 component. ===Camera module=== Camera modules can't be directly connected to APF signals, but need a breakout board. Moreover, since they need to be put in a different place than the processing board, they need a connector on the main board. ===FPGA connectors for servo PWM=== If we chose servo as the actuators for the airplane ailerons, we will need a PWM signal to control them. The FPGA of Armadeus boards is able to control lots of them as explained [http://www.fpga4fun.com/RCServos.html here], but we need to add the connectors to the main board. 61dd730d2a56f03b7eec964e032b928966880062 6 20 87 2010-10-23T14:17:54Z Vincent 1 The datasheet of the DS1374 I2C 32-bit binary counter watchdog RTC with trickle charger and reset input/output. wikitext text/x-wiki The datasheet of the DS1374 I2C 32-bit binary counter watchdog RTC with trickle charger and reset input/output. b846e6d6b75f2344910f54240d2828a0385a6fcf 12 4 88 7 2010-10-23T14:22:36Z Vincent 1 /* Internal links */ wikitext text/x-wiki You can format your text using wiki markup. This consists of normal characters like asterisks, single quotes or equation marks which have a special function in the wiki, sometimes depending on their position. For example, to format a word in ''italic'', you include it in two single quotes like <code><nowiki>''this''</nowiki></code>. == Text formatting markup == {| class="wikitable" ! Description !! You type !! 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This is supposed to be end user help is it not? --> setting. }} {{admin tip|tip= To remove the “external link icons“ from next to each of the external links, add the following to the page located at <code>MediaWiki:Monobook.css</code> on your wiki. <source lang="css"> #bodyContent a.external, #bodyContent a[href ^="gopher://"] { background: none; padding-right: 0; } </source> }} === How to avoid auto-links === By default, when you write a URL as is, it will be transformed to an external link. To avoid that effect, put the URL between <code>&lt;nowiki&gt;</code> tags as in: <nowiki><nowiki>http://mediawiki.org</nowiki></nowiki> [[Category:Help|Formatting]] 1ec117c8f657181fb114eef6674012ce75b048ba 6 21 100 2010-11-17T08:43:54Z Vincent 1 Cut of the injector plate of the RD-107 engine wikitext text/x-wiki Cut of the injector plate of the RD-107 engine 172271a433cf40c486047014f262044838a804ca 0 17 101 97 2010-11-17T09:45:23Z Vincent 1 /* Rocket Engine */ adding cut drawings for RD-107 injector wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |Yes, with Nitrogen |No |No |Yes, with Helium |- |'''Fuel pump''' |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |'''Injector''' |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |Copper sheet inside, steel outside? |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic |? | | |Electric |- |'''Provided by''' |Engine's turbopumps |? | | |? |- |'''Actuator''' |Six hydraulic servoactuators |? |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Traditionnaly, only turbo pumps have been able to feed the engine at a large enough rate. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. ==Cooling== Regenerative cooling is most widely used in rocket engines. Few of them however use other ways, like ablatively cooling carbon fiber composite in SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine, or radiative cooling in the Merlin Vacuum nozzle (still regenerative for the chamber). 7fff738ec19a5abaad0d8ff57145594f07bf804b 102 101 2010-11-17T20:21:05Z Vincent 1 Soyuz fixes wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |'''Fuel pump''' |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |'''Injector''' |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Traditionnaly, only turbo pumps have been able to feed the engine at a large enough rate. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. ==Cooling== Regenerative cooling is most widely used in rocket engines. Few of them however use other ways, like ablatively cooling carbon fiber composite in SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine, or radiative cooling in the Merlin Vacuum nozzle (still regenerative for the chamber). ada308c3cc26cead67dd0a3f2bcded69d2fd9450 103 102 2010-11-18T00:13:17Z Vincent 1 /* Cooling */ begin of the cooling part wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |'''Fuel pump''' |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |'''Injector''' |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Traditionnaly, only turbo pumps have been able to feed the engine at a large enough rate. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidant, to cool the chamber walls before injecting those propellants into the chamber. The cooler flows into a series of pipes or crafting in the external or intermediate wall of the engine, around the nozzle, around the chamber, or around both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). aacaa3cdebc6042ae2d6b87620286c095eee29ae 104 103 2010-11-18T00:52:46Z Vincent 1 /* Cooling */ lox as coolant wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |'''Fuel pump''' |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |'''Injector''' |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Traditionnaly, only turbo pumps have been able to feed the engine at a large enough rate. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, and not fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine cutoff. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotory Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Moreover, LOX and cheap fuels are readily available. d41ef5b3d60d3d36603b763b307fd6689f61eec1 108 104 2010-11-20T00:39:23Z Vincent 1 /* Cooling */ adding 'cooling for LOX/E85 engine' title wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |'''Fuel pump''' |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |'''Injector''' |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Traditionnaly, only turbo pumps have been able to feed the engine at a large enough rate. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, and not fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine cutoff. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotory Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Moreover, LOX and cheap fuels are readily available. 8b7ee2390dc1c58beae6fbcb1976b4576b03b246 111 108 2010-12-24T17:41:47Z Vincent 1 /* Pumps and tank pressurization */ tank link wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |'''Fuel pump''' |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |'''Injector''' |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, and not fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine cutoff. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotory Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Moreover, LOX and cheap fuels are readily available. 594216f0cbdd76bcf73a25fcf1a86def7b9f59f8 114 111 2010-12-24T18:02:12Z Vincent 1 pressurization section wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of existing rocket engines. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |'''Tank pressurization''' |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |'''Fuel pump''' |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |'''Cooling''' |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |'''Injector''' |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, and not fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine cutoff. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotory Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Moreover, LOX and cheap fuels are readily available. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. 033d1bf604baed831995b65fe26ea9bb9a502432 120 114 2011-01-03T21:36:24Z Vincent 1 Links in the table to anchors, adding injector section. wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, and not fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine cutoff. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotory Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Moreover, LOX and cheap fuels are readily available. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, and are able to deep throttle up to 1:35. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. aee0e7c22196dbd6067e38c01e6a0ccd81fd19e1 121 120 2011-01-03T21:43:52Z Vincent 1 /* Injectors */ drawbacks. wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, and not fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine cutoff. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotory Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Moreover, LOX and cheap fuels are readily available. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. The Lunar Module Descent Engine is the most famous pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, and are able to deep throttle up to 1:35. It has never been seen that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. 37948d2a8cbfe994f88f7cf5606d97e9bd13a479 122 121 2011-01-03T21:53:35Z Vincent 1 /* Injectors */ SpaceX and more info wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, and not fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine cutoff. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotory Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Moreover, LOX and cheap fuels are readily available. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. From SpaceX's experience, it seems that regenerative cooling is sufficient to sustain the combustion temperature hitting the walls. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. ac0543fae8ec5bf63146a6e2a4508de92f173cdd 0 7 105 17 2010-11-20T00:06:14Z Vincent 1 title wikitext text/x-wiki =N-Prize= With the creation of some competitions for private-funded companies/projects in the space domain a few years ago, Dr. Paul H. Dear thought about an other competition open to imaginative people. The idea was [http://www.halfbakery.com/idea/N-Prize brainstormed] and the competition was finally created, along with strict [http://www.n-prize.com rules]. The goal of the competition is to launch from earth surface a very small object into orbit. This satellite must not exceed 20 grams, and the difficult thing is to be able to make this orbital insertion for a cost less than £1,000. There are two sets of rules: the launch can be single use (Single-Spent-to-orbit), like a rocket, and thus using the £1,000 for the full price of the launch vehicle and the satellite. The other way is a reusable launch vehicule (RV) to some extend, for example recovering the first stage of a rocket or a Helium balloon. Everything that is recovered and reusable doesn't count into the £1,000. The pushed-into-orbit [10-20[g satellite has to be tracked by entrants on their own. To be validated, proof must be established that the satellite has made at least 9 orbits before re-entry. b3906f20ac1d863b01c68e16b1be530bb8ebbf42 4 22 106 2010-11-20T00:07:59Z Vincent 1 Redirecting to [[N-Prize]] wikitext text/x-wiki #REDIRECT [[N-Prize]] 26349981933f7a8be9cbff7bdb57bc6cedca4aa5 8 2 107 84 2010-11-20T00:13:10Z Vincent 1 adding rocket engine link wikitext text/x-wiki * navigation ** mainpage|mainpage ** Resources|Resources ** recentchanges-url|recentchanges ** randompage-url|randompage * sections ** EmbeddedRocketComputer|Computer ** Build_a_cheap_turbofan|Plane engine ** RocketEngines|Rocket engine ** Testing|Testing and validation 3cde20a1dd597668d5fc7f08ecb808361e649a2d 0 1 109 71 2010-11-20T00:44:06Z Vincent 1 /* Fuel */ more info on E85 and link to cooling in rocket engines wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. For the cooling, I see [http://en.wikipedia.org/wiki/Regenerative_cooling_(rocket) regenerative cooling] as the only option. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. ba31185cd6274dc2ab596544500e43a2c9914802 110 109 2010-11-20T00:46:27Z Vincent 1 /* Engine */ removing cooling sentence wikitext text/x-wiki =N-Prize reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 11d2546b367fb7bdf2f8717de8938eb1df05d4db 125 110 2011-01-24T23:34:33Z Vincent 1 /* N-Prize reflections */ wikitext text/x-wiki =N-Prize and low-cost space access reflections= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 127b8e78dcaad4eb228701ce35ec028416d8ac5d 6 23 112 2010-12-24T17:45:33Z Vincent 1 Center unit of the redstone launcher, basically made of the propellant tank. From http://www.myarmyredstonedays.com/Photos/page8/shell_03.html . wikitext text/x-wiki Center unit of the redstone launcher, basically made of the propellant tank. From http://www.myarmyredstonedays.com/Photos/page8/shell_03.html . aabd99087fe771693ab88a264623f2b4c2c64c2a 0 24 113 2010-12-24T17:48:51Z Vincent 1 Page creation and adding redstone tank image. wikitext text/x-wiki =Rocket Fuel tanks= In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The redstone rocket for example used a single tank with an internal separation, as we can see below. [[Image:Restone_tank.jpg|250px]] 12059ca551f4301b9d5f7114709ee316823ce15f 115 113 2010-12-24T18:51:06Z Vincent 1 /* Rocket Fuel tanks */ sloshing wikitext text/x-wiki =Rocket Fuel tanks= In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The redstone rocket for example used a single tank with an internal separation, as we can see below. [[Image:Restone_tank.jpg|center|250px]] ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. 1592bcdd3c73992cfe9105ce2aa503e8eb53cede 6 25 116 2010-12-29T03:58:13Z Vincent 1 BOOK: HOW to DESIGN, BUILD and TEST SMALL LIQUID-FUEL ROCKET ENGINES Backup archive from http://www.risacher.org/rocket/ wikitext text/x-wiki BOOK: HOW to DESIGN, BUILD and TEST SMALL LIQUID-FUEL ROCKET ENGINES Backup archive from http://www.risacher.org/rocket/ 5eac731f51d353be5afe33458eceade1385de9d8 118 116 2010-12-29T04:09:51Z Vincent 1 wikitext text/x-wiki BOOK: HOW to DESIGN, BUILD and TEST SMALL LIQUID-FUEL ROCKET ENGINES Backup archive from http://www.risacher.org/rocket/ Rocketlab / China lake, Calif. 1967. 87d02b3a5e245c1f605a7b63f6327bac48443fe4 6 26 117 2010-12-29T04:01:22Z Vincent 1 TRW pintle engine heritage and performance characteristics Gordon A. Dressler, J. Martin Bauer. 2000. TRW copyright. wikitext text/x-wiki TRW pintle engine heritage and performance characteristics Gordon A. Dressler, J. Martin Bauer. 2000. TRW copyright. f77ecd30bfac043f95098d1e08071a55029c9626 0 16 119 99 2010-12-29T04:14:00Z Vincent 1 /* Rocket engines */ adding book 'how to design, build and test... wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Web pages== * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Books (online links)== ===Multi-domain=== * [http://books.google.fr/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.fr/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.fr/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. ====Specific parts:==== * [http://books.google.fr/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.fr/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.fr/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.fr/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.fr/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.fr/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.fr/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. 51ad37d9c022741036472f885f4e8a02478b8388 128 119 2011-02-02T10:58:45Z Vincent 1 /* Videos (youtube links) */ turbines channel wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Web pages== * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Books (online links)== ===Multi-domain=== * [http://books.google.fr/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.fr/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.fr/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. ====Specific parts:==== * [http://books.google.fr/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.fr/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.fr/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.fr/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.fr/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.fr/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.fr/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. 7f59ec1e93cce9c58b3b22cfaf40d49368a4f9c8 0 11 123 59 2011-01-09T02:55:50Z Vincent 1 new subsections for the "our design" part, and spelling corrections everywhere wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== A main optimization to reduce cost and complexity would be to create a simple design of parts creating the internal volumes of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [[http://en.wikipedia.org/wiki/Lathe lathe]] ([http://en.wikipedia.org/wiki/Turning turning]). [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Our design== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C... ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor and stator=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. Also, the beginning of the blade (out of the shaft) has to be linear and the insert has to be parallel to the shaft. It imposes the blades to be twisted between the inner part and the large outer part. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. e5852551300a1cc1f01cc5576754f1d5cb0aabd2 124 123 2011-01-09T03:00:05Z Vincent 1 /* Shaped core or shaped shaft? */ fixes wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Maybe the only machined shaft design is not possible in real-life for some reason, and it has to be studied. [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Our design== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C... ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor and stator=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. Also, the beginning of the blade (out of the shaft) has to be linear and the insert has to be parallel to the shaft. It imposes the blades to be twisted between the inner part and the large outer part. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ba55c1fa6ca8f236d359615cfe42fad38f500a1d 126 124 2011-01-24T23:37:03Z Vincent 1 /* Our design */ wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Maybe the only machined shaft design is not possible in real-life for some reason, and it has to be studied. [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Our design== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C... ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor and stator=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. Also, the beginning of the blade (out of the shaft) has to be linear and the insert has to be parallel to the shaft. It imposes the blades to be twisted between the inner part and the large outer part. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. fac091bc9e851723da01573407931b30919de50c 127 126 2011-02-02T10:57:41Z Vincent 1 /* General principles */ youtube channel to turbines wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Maybe the only machined shaft design is not possible in real-life for some reason, and it has to be studied. [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Our design== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C... ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor and stator=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. Also, the beginning of the blade (out of the shaft) has to be linear and the insert has to be parallel to the shaft. It imposes the blades to be twisted between the inner part and the large outer part. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. d78396dd3effd0e0bd60beaf2f6f01562e667f7a 129 127 2011-02-02T13:54:15Z Vincent 1 /* Shaped core or shaped shaft? */ hollow stage design wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, and it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. The issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Our design== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C... ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor and stator=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. Also, the beginning of the blade (out of the shaft) has to be linear and the insert has to be parallel to the shaft. It imposes the blades to be twisted between the inner part and the large outer part. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. 057727232e037fafa90e134cd9b94501b65959ad 132 129 2011-02-22T01:54:54Z Vincent 1 creating new section for proposed design wikitext text/x-wiki =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, and it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. The issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Design considerations== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C... ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor and stator=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. Also, the beginning of the blade (out of the shaft) has to be linear and the insert has to be parallel to the shaft. It imposes the blades to be twisted between the inner part and the large outer part. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ==Our Design propositions== From the different concerns expressed above, we propose a new design for a low-cost turbofan. Several pages have been created to explain each subsystem and parts manufacturability. * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. 68c621ff407db57876a8c524577f418f26f18820 142 132 2011-02-22T23:27:28Z Vincent 1 Adding category and top text wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, and it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. The issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Design considerations== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C... ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor and stator=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. Also, the beginning of the blade (out of the shaft) has to be linear and the insert has to be parallel to the shaft. It imposes the blades to be twisted between the inner part and the large outer part. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ==Our Design propositions== From the different concerns expressed above, we propose a new design for a low-cost turbofan. Several pages have been created to explain each subsystem and parts manufacturability. * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. [[Category:Turbofan|Index]] be284f0fc72712bea60fca9bd5183c3c94d52cf4 146 142 2011-02-22T23:37:10Z Vincent 1 wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, and it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. The issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Design considerations== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C... ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor and stator=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. Also, the beginning of the blade (out of the shaft) has to be linear and the insert has to be parallel to the shaft. It imposes the blades to be twisted between the inner part and the large outer part. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ==Our Design propositions== From the different concerns expressed above, we propose a new design for a low-cost turbofan. Several pages have been created to explain each subsystem and parts manufacturability. * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. [[Category:Turbofan]] 231d7213d4868b5dd8c00ff6590c2910a002ade5 147 146 2011-02-23T00:07:09Z Vincent 1 /* Fixing blades to rotor and stator */ wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, and it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. The issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Design considerations== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C... ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades, and maybe a simple design similar to the one above, but based on only one squared holder is enough. A completely different design of the mounting would be to use spacer parts between each blade, that would create the whole compressor stage fixation when all blades and spacer are put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more eavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages. Indeed, higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft. Creating a part that can be duplicated to form a cylinder, but with a non-parallel to the axis separation poses design and crafting issues. An alternative design would be to have the blades orthogonally free (axially too) and fix them on the sides using a key of some sort. The blades mounting ring and the blades would have a notch in the same place, in which a key would place to hold everything together. ===Fixing blades to stator=== To be studied. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ==Our Design propositions== From the different concerns expressed above, we propose a new design for a low-cost turbofan. Several pages have been created to explain each subsystem and parts manufacturability. * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. [[Category:Turbofan]] fafda305649b6c4f0a1d05325708f437e1e265c0 4 27 130 2011-02-21T22:08:26Z Vincent 1 Created page with "All material on this wiki is protected by the [[http://www.gnu.org/copyleft/fdl.html GNU Free Documentation License]], meaning it's free of use for non-commercial purposes, and a…" wikitext text/x-wiki All material on this wiki is protected by the [[http://www.gnu.org/copyleft/fdl.html GNU Free Documentation License]], meaning it's free of use for non-commercial purposes, and apply in more [[http://en.wikipedia.org/wiki/GFDL_1.3 complicated ways]] to commercial uses. The idea is to share knowledge and make human being move forward, not to make money and move backwards because of secrets. bf72485ccf327a35fd099c43ff6ca6f50fae55b2 131 130 2011-02-21T22:08:45Z Vincent 1 wikitext text/x-wiki All material on this wiki is protected by the [http://www.gnu.org/copyleft/fdl.html GNU Free Documentation License], meaning it's free of use for non-commercial purposes, and apply in more [http://en.wikipedia.org/wiki/GFDL_1.3 complicated ways] to commercial uses. The idea is to share knowledge and make human being move forward, not to make money and move backwards because of secrets. c0d066f907cf7a603618e30b8fdd5dae5e7abd79 6 28 133 2011-02-22T01:58:39Z Vincent 1 Exploded view of a simple compressor blade fixation design. Between each level, a conic piece like the yellow on this picture is placed and plugs onto a new gray piece like the one in the background, which supports the blades. wikitext text/x-wiki Exploded view of a simple compressor blade fixation design. Between each level, a conic piece like the yellow on this picture is placed and plugs onto a new gray piece like the one in the background, which supports the blades. 85fae216396702919d942d7a307adb1448f3d227 137 133 2011-02-22T02:18:13Z Vincent 1 wikitext text/x-wiki Exploded view of a simple compressor blade fixation design. Between each level, a conic piece like the yellow on this picture is placed and plugs onto a new gray piece like the one in the background, which supports the blades. PARTS ARE NOT TO SCALE. b5644d7c33227b53e4abde156306c8e02e4fad10 6 29 134 2011-02-22T02:03:52Z Vincent 1 Exploded view of a simple compressor blade fixation design. The main shaft (orange) has one spline at least, allowing the blade support and the conic parts to be coupled to the shaft. wikitext text/x-wiki Exploded view of a simple compressor blade fixation design. The main shaft (orange) has one spline at least, allowing the blade support and the conic parts to be coupled to the shaft. 0786216481edb3a8352412ce90c1b4cd1bac315f 138 134 2011-02-22T02:18:32Z Vincent 1 wikitext text/x-wiki Exploded view of a simple compressor blade fixation design. The main shaft (orange) has one spline at least, allowing the blade support and the conic parts to be coupled to the shaft. PARTS ARE NOT TO SCALE. 0cf8b84ab81bd4a737820e278c7da06f660766c4 0 30 135 2011-02-22T02:10:13Z Vincent 1 Creating page with first two images. wikitext text/x-wiki =Compressor design= ==Compressor blades mounting== [[File:Interblade1.jpg|600px|center]] [[File:Interblade2.jpg|600px|center]] ==Compressor blade close-up and manufacturing== ''Coming soon.'' 28f02040aa879cc068845d675b0333998b186408 136 135 2011-02-22T02:15:54Z Vincent 1 /* Compressor blades mounting */ text wikitext text/x-wiki =Compressor design= ==Compressor blades mounting== For the first compressor stage and the fan, the design is simplified by putting the roots of the blades parallel to the shaft. On stages 2 and 3, blades are not long enough, and to finish at the correct angle and not be twisted too much, they need to start at a non-zero angle. That brings problems to the design of the blade socket and fixing. The two pictures below propose a solution to this problem: blades' roots remain linear, the milling of the mounting is linear too, while it's a round part. This makes a strange effect, but allows the blades to be properly fixed on their sides, and if not enough, by a keyed design. [[File:Interblade1.jpg|600px|center]] [[File:Interblade2.jpg|600px|center]] ==Compressor blade close-up and manufacturing== ''Coming soon.'' 0ea1024c789dfd5bec2bbde9508e087e52bd9795 139 136 2011-02-22T02:22:14Z Vincent 1 stator/rotor wikitext text/x-wiki =Compressor design= ==Overall view and notes== ==Rotor design== ===Compressor blades mounting=== For the first compressor stage and the fan, the design is simplified by putting the roots of the blades parallel to the shaft. On stages 2 and 3, blades are not long enough, and to finish at the correct angle and not be twisted too much, they need to start at a non-zero angle. That brings problems to the design of the blade socket and fixing. The two pictures below propose a solution to this problem: blades' roots remain linear, the milling of the mounting is linear too, while it's a round part. This makes a strange effect, but allows the blades to be properly fixed on their sides, and if not enough, by a keyed design. [[File:Interblade1.jpg|600px|center]] [[File:Interblade2.jpg|600px|center]] ===2nd and 3rd stage's compressor blade close-up and manufacturing=== ''Coming soon.'' ==Stator design== Even more complicated. 7692c1e53c7c29ab07fb29a1515e03d713a37df6 140 139 2011-02-22T02:22:55Z Vincent 1 /* Compressor design */ wikitext text/x-wiki =Compressor design= ==Rotor design== ===Compressor blades mounting=== For the first compressor stage and the fan, the design is simplified by putting the roots of the blades parallel to the shaft. On stages 2 and 3, blades are not long enough, and to finish at the correct angle and not be twisted too much, they need to start at a non-zero angle. That brings problems to the design of the blade socket and fixing. The two pictures below propose a solution to this problem: blades' roots remain linear, the milling of the mounting is linear too, while it's a round part. This makes a strange effect, but allows the blades to be properly fixed on their sides, and if not enough, by a keyed design. [[File:Interblade1.jpg|600px|center]] [[File:Interblade2.jpg|600px|center]] ===2nd and 3rd stage's compressor blade close-up and manufacturing=== ''Coming soon.'' ==Stator design== Even more complicated. 73fb791d614a76e415649e28cc285cb25972ee94 141 140 2011-02-22T02:27:29Z Vincent 1 /* Compressor design */ wikitext text/x-wiki =Compressor design= Real-world engines can have nearly 20 compression stages. We will first build a 3-stage compressor, in a single-spool engine to assess how it can be scaled to a full power engine, or if it will require more stages (= quite costly). ==Rotor design== Special care has to be given to the rotor, since it will spin at very high rotation speeds (not calculated yet). A blade mounting failure will likely cause the loss of the aircraft, given the constraints. ===Compressor blades mounting=== For the first compressor stage and the fan, the design is simplified by putting the roots of the blades parallel to the shaft. On stages 2 and 3, blades are not long enough, and to finish at the correct angle and not be twisted too much, they need to start at a non-zero angle. That brings problems to the design of the blade socket and fixing. The two pictures below propose a solution to this problem: blades' roots remain linear, the milling of the mounting is linear too, while it's a round part. This makes a strange effect, but allows the blades to be properly fixed on their sides, and if not enough, by a keyed design. [[File:Interblade1.jpg|600px|center]] [[File:Interblade2.jpg|600px|center]] ===2nd and 3rd stage's compressor blade close-up and manufacturing=== ''Coming soon.'' ==Stator design== Even more complicated. 857eb6b795bc45401625c321dac034c074c7c262 143 141 2011-02-22T23:28:22Z Vincent 1 adding category wikitext text/x-wiki =Compressor design= Real-world engines can have nearly 20 compression stages. We will first build a 3-stage compressor, in a single-spool engine to assess how it can be scaled to a full power engine, or if it will require more stages (= quite costly). ==Rotor design== Special care has to be given to the rotor, since it will spin at very high rotation speeds (not calculated yet). A blade mounting failure will likely cause the loss of the aircraft, given the constraints. ===Compressor blades mounting=== For the first compressor stage and the fan, the design is simplified by putting the roots of the blades parallel to the shaft. On stages 2 and 3, blades are not long enough, and to finish at the correct angle and not be twisted too much, they need to start at a non-zero angle. That brings problems to the design of the blade socket and fixing. The two pictures below propose a solution to this problem: blades' roots remain linear, the milling of the mounting is linear too, while it's a round part. This makes a strange effect, but allows the blades to be properly fixed on their sides, and if not enough, by a keyed design. [[File:Interblade1.jpg|600px|center]] [[File:Interblade2.jpg|600px|center]] ===2nd and 3rd stage's compressor blade close-up and manufacturing=== ''Coming soon.'' ==Stator design== Even more complicated. [[Category:Turbofan]] fbef72ee8287b36f9cf175b090d36b68d2d3777e 144 143 2011-02-22T23:33:31Z Vincent 1 wikitext text/x-wiki =Compressor design= Real-world engines can have nearly 20 compression stages. We will first build a 3-stage compressor, in a single-spool engine to assess how it can be scaled to a full power engine, or if it will require more stages (= quite costly). ==Rotor design== Special care has to be given to the rotor, since it will spin at very high rotation speeds (not calculated yet). A blade mounting failure will likely cause the loss of the aircraft, given the constraints. ===Compressor blades mounting=== For the first compressor stage and the fan, the design is simplified by putting the roots of the blades parallel to the shaft. On stages 2 and 3, blades are not long enough, and to finish at the correct angle and not be twisted too much, they need to start at a non-zero angle. That brings problems to the design of the blade socket and fixing. The two pictures below propose a solution to this problem: blades' roots remain linear, the milling of the mounting is linear too, while it's a round part. This makes a strange effect, but allows the blades to be properly fixed on their sides, and if not enough, by a keyed design. [[File:Interblade1.jpg|600px|center]] [[File:Interblade2.jpg|600px|center]] ===2nd and 3rd stage's compressor blade close-up and manufacturing=== ''Coming soon.'' ==Stator design== Even more complicated. [[Category:Turbofan|Compressor]] 5af2a6b003dac6e4972616bbd1212396d424db69 148 144 2011-02-23T00:12:25Z Vincent 1 blade page link wikitext text/x-wiki =Compressor design= Real-world engines can have nearly 20 compression stages. We will first build a 3-stage compressor, in a single-spool engine to assess how it can be scaled to a full power engine, or if it will require more stages (= quite costly). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. ==Rotor design== Special care has to be given to the rotor, since it will spin at very high rotation speeds (not calculated yet). A blade mounting failure will likely cause the loss of the aircraft, given the constraints. ===Compressor blades mounting=== For the first compressor stage and the fan, the design is simplified by putting the roots of the blades parallel to the shaft. On stages 2 and 3, blades are not long enough, and to finish at the correct angle and not be twisted too much, they need to start at a non-zero angle. That brings problems to the design of the blade socket and fixing. The two pictures below propose a solution to this problem: blades' roots remain linear, the milling of the mounting is linear too, while it's a round part. This makes a strange effect, but allows the blades to be properly fixed on their sides, and if not enough, by a keyed design. [[File:Interblade1.jpg|600px|center]] [[File:Interblade2.jpg|600px|center]] ===2nd and 3rd stage's compressor blade close-up and manufacturing=== ''Coming soon.'' ==Stator design== Even more complicated. [[Category:Turbofan|Compressor]] bede5f6b5b446e7b69ac9152a207e6a978e16603 14 31 145 2011-02-22T23:35:29Z Vincent 1 Created page with "The Turbofan category groups pages related to our turbofan's design proposition. On those pages you can find explanations for every part of the engine, schematics, manufacturing …" wikitext text/x-wiki The Turbofan category groups pages related to our turbofan's design proposition. On those pages you can find explanations for every part of the engine, schematics, manufacturing details, and so on. 551ccede11b84ea1026decc846f47472f043b53a 0 32 149 2011-02-23T00:38:55Z Vincent 1 Created page with "=Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. The root/fixation…" wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. ==Blade designs for efficient air flow== Rotor and stator blades have to be carefully designed, since they provide the turbine all its power. Leaks (free air paths) have to be minimized. Swirls have to be avoided in the compressor and turbine for several stages to work. For that reason, the stator deflects air in the opposite direction than the rotor, allowing the next stage to perform as if it receives untouched air, or even better oriented in the most efficient direction for the rotor. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing] is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing Hot isostatic pressing] possibly too. I believe that a hot forging press can be done cheaply considering the size of the blades. Work-hardening and head-treating have to be studied, and depend on the crafting method used in the first place. The metal used for the blades may be an aluminum alloy for the compressor, and a steel or nickel-rich alloy for the turbine because of heat. Don't forget that the blade insert will have to be milled at some point. [[Category:Turbofan|Compressor]] da3576b4eab525b3484c68dc372559adff126377 150 149 2011-02-23T00:53:25Z Vincent 1 /* Manufacturing propositions */ wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. ==Blade designs for efficient air flow== Rotor and stator blades have to be carefully designed, since they provide the turbine all its power. Leaks (free air paths) have to be minimized. Swirls have to be avoided in the compressor and turbine for several stages to work. For that reason, the stator deflects air in the opposite direction than the rotor, allowing the next stage to perform as if it receives untouched air, or even better oriented in the most efficient direction for the rotor. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing] is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing hot isostatic pressing] possibly too, as explained on the ''How are made turbine blades'' [http://www.youtube.com/watch?v=vN3_Wkyl5PQ video]. I believe that a hot forging press can be done cheaply considering the small size of our blades. For the main fan, it thus may not be used. Work-hardening and head-treating have to be studied, and depend on the crafting method used in the first place. The metal used for the blades may be an aluminum alloy for the compressor, and a steel or nickel-rich alloy for the turbine because of heat. Don't forget that the blade insert will have to be milled at some point. [[Category:Turbofan|Compressor]] a94fb17388cbfe8917e338770c3ae0cf284eb19d 0 32 151 150 2011-02-23T00:54:05Z Vincent 1 /* Manufacturing propositions */ wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. ==Blade designs for efficient air flow== Rotor and stator blades have to be carefully designed, since they provide the turbine all its power. Leaks (free air paths) have to be minimized. Swirls have to be avoided in the compressor and turbine for several stages to work. For that reason, the stator deflects air in the opposite direction than the rotor, allowing the next stage to perform as if it receives untouched air, or even better oriented in the most efficient direction for the rotor. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing] is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing hot isostatic pressing] possibly too, as explained on the ''How are made turbine blades'' [http://www.youtube.com/watch?v=vN3_Wkyl5PQ video]. I believe that a hot forging press can be done cheaply considering the small size of our blades. For the main fan, it thus may not be used. Work-hardening and head-treating have to be studied, and depend on the crafting method used in the first place. The metal used for the blades may be an aluminum alloy for the compressor, and a steel or nickel-rich alloy for the turbine because of heat. Don't forget that the blade insert will have to be milled at some point. [[Category:Turbofan|Blades]] 2732e6aae3405bdfdb58898e262222525deba142 154 151 2011-02-23T01:14:37Z Vincent 1 copying text from the build a cheap turbofan page wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. ==Blade designs for efficient air flow== Blades have to be carefully designed, since they provide the turbine all its power. Stages are not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing] is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing hot isostatic pressing] possibly too, as explained on the ''How are made turbine blades'' [http://www.youtube.com/watch?v=vN3_Wkyl5PQ video]. I believe that a hot forging press can be done cheaply considering the small size of our blades. For the main fan, it thus may not be used. Work-hardening and head-treating have to be studied, and depend on the crafting method used in the first place. The high-pressure turbine blades have to face very high temperature and pressure. On real engines, they are made of titanium and nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. Since the required lifetime is lower in our case, we may achieve a working engine with cheaper metals, like steel or nickel-rich alloy for the turbine blades. For the compressor blades, aluminum alloys are probably be a good solution. Don't forget that the blade insert will have to be milled at some point. [[Category:Turbofan|Blades]] 285aa082e23ee4be91118a7483511ced29192309 158 154 2011-02-23T22:54:35Z Vincent 1 /* Manufacturing propositions */ adding picture wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. ==Blade designs for efficient air flow== Blades have to be carefully designed, since they provide the turbine all its power. Stages are not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing] is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing hot isostatic pressing] possibly too, as explained on the ''How are made turbine blades'' [http://www.youtube.com/watch?v=vN3_Wkyl5PQ video]. I believe that a hot forging press can be done cheaply considering the small size of our blades. For the main fan, it thus may not be used. [[File:Blade.jpg|600px|center]] Work-hardening and head-treating have to be studied, and depend on the crafting method used in the first place. The high-pressure turbine blades have to face very high temperature and pressure. On real engines, they are made of titanium and nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. Since the required lifetime is lower in our case, we may achieve a working engine with cheaper metals, like steel or nickel-rich alloy for the turbine blades. For the compressor blades, aluminum alloys are probably be a good solution. Don't forget that the blade insert will have to be milled at some point. [[Category:Turbofan|Blades]] 2ed375ffe7de48bc7ee6afd4d55b78e25c9722e7 198 158 2011-05-04T23:46:46Z Vincent 1 /* Blade design and manufacturing */ new outline of the page and more information on everything. wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. ==Blade fixation== The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. In summary, real engines use a fir tree shape to hold the blade centrifugally, while keeping them free axially. They are generally fixed using a locking screw. This design allows for easy replacement of damaged blades, but is quite complex to manufacture, and requires lots of parts. Since our engine will be smaller, using fixing screws may not be feasible, and blades are subject to less centrifugal stress. Simpler blade fixation mechanism should be relevant. ==Blade designs for efficient air flow== Blades have to be carefully designed, because the overall efficiency of the engine largely depends on their design. '''Stages.''' They are not only made of blades on the rotor, but also blades on the stator, generally called ''vanes''. A '''stage''' is then a pair of a rotor stage and a stator stage, in this order. The stator prevents a rotating air flow to form inside the engine (swirl), driven by the action of the rotor blades. Stator vanes redirect the airflow in the more appropriate direction for the next rotor stage. They increase the energy of the gas ([http://en.wikipedia.org/wiki/Enthalpy enthalpy]) by removing the swirling effect that impairs it. '''Blade shape.''' Most basic design of a fan has flat-shaped blades. Twisted blades with a flat section are an improvement, taking into account the difference in apparent airflow velocity and torque all along the blade. Next step is to have a non-flat section, but an airfoil section. This has been proved to be the only design meeting efficiency requirements of the turbine engines, in 192X by XXX. Finally, modern engines are designed with curved edges for the fan, for optimal known efficiency as well as for noise reduction. '''Rotor/stator gaps.''' Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. Design of stages is linked to the energy the blades have to give (compressor and fan) or take (turbine) to the air flow. To better define and understand that energy, we will use standard [http://en.wikipedia.org/wiki/State_function thermodynamic parameters] of gas, a.k.a state variables of a gas, on which are based quantities like the enthalpy: temperature, volume, pressure. We will also use the velocity because the actual work given by a turbofan engine is related to the mass flow rate of the gas expelled by the engine, which relates to velocity of this gas and the state variables. ===Gas variables: temperature, pressure, velocity=== ===Fan design=== Main goal of the fan is to increase the mass flow rate, mainly by increasing the velocity. The mass flow is linked to the area of the fan blades and the angular speed of the fan. To increase the velocity, simplest way is to reduce volume. The fan duct will thus have to act as a compressor on the aft-end. On the front-end, it is generally designed as an expander, to increase the pressure, allowing more efficient work on the air flow. ===Compressor design=== Main goal of the compressor is to increase pressure. Due to friction of the gas on the blades and guide vanes mainly, temperature is increased too. Thus, volume has to decrease, intake area will be greater than compressor discharge area. ===Turbine design=== Main goal of the turbine is to extract energy from the hot and fast gas discharged by the combustion into mechanical (rotational) work. Pressure and temperature may remain constant through the turbine, and high velocity and pressure will provide better efficiency ''[to be verified]''. ==Mechanical constraints== Blades on all three parts of a turbofan engine undergo heavy mechanical constraints due to high rpm achieved by the rotor, the high temperature in the turbine section and non-negligible temperature in the end of the compressor section, and the high pressure of the gas on which work is performed. Fan blades are not made of plain metal in real engines. In the eighties, they were made in honeycomb composite sandwich material, they are now made in triangular sandwich structure. Compressor blades are made of titanium alloys, providing high strength and rigidity at these temperatures. Turbine blades are made of nickel alloys, better sustaining the high temperature, and still at higher strength than steel. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing] is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing hot isostatic pressing] possibly too, as explained on the ''How are made turbine blades'' [http://www.youtube.com/watch?v=vN3_Wkyl5PQ video]. I believe that a hot forging press can be done cheaply considering the small size of our blades. For the main fan, it thus may not be used. [[File:Blade.jpg|600px|center]] ''The above picture lacks the airfoil section of the blade, because the CAD software we use doesn't currently support it.'' Work-hardening and head-treating have to be studied, and depend on the crafting method used in the first place. The high-pressure turbine blades have to face very high temperature and pressure. On real engines, they are made of titanium and nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. Since the required lifetime is lower in our case, we may achieve a working engine with cheaper metals, like steel or nickel-rich alloy for the turbine blades. For the compressor blades, aluminum alloys are probably be a good solution. Don't forget that the blade insert will have to be milled at some point. [[Category:Turbofan|Blades]] 3ae8eefdfaa3dfe73821d9df5d2552bb30b34510 199 198 2011-05-05T00:06:08Z Vincent 1 /* Mechanical constraints */ HST Schuebeler link wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. ==Blade fixation== The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. In summary, real engines use a fir tree shape to hold the blade centrifugally, while keeping them free axially. They are generally fixed using a locking screw. This design allows for easy replacement of damaged blades, but is quite complex to manufacture, and requires lots of parts. Since our engine will be smaller, using fixing screws may not be feasible, and blades are subject to less centrifugal stress. Simpler blade fixation mechanism should be relevant. ==Blade designs for efficient air flow== Blades have to be carefully designed, because the overall efficiency of the engine largely depends on their design. '''Stages.''' They are not only made of blades on the rotor, but also blades on the stator, generally called ''vanes''. A '''stage''' is then a pair of a rotor stage and a stator stage, in this order. The stator prevents a rotating air flow to form inside the engine (swirl), driven by the action of the rotor blades. Stator vanes redirect the airflow in the more appropriate direction for the next rotor stage. They increase the energy of the gas ([http://en.wikipedia.org/wiki/Enthalpy enthalpy]) by removing the swirling effect that impairs it. '''Blade shape.''' Most basic design of a fan has flat-shaped blades. Twisted blades with a flat section are an improvement, taking into account the difference in apparent airflow velocity and torque all along the blade. Next step is to have a non-flat section, but an airfoil section. This has been proved to be the only design meeting efficiency requirements of the turbine engines, in 192X by XXX. Finally, modern engines are designed with curved edges for the fan, for optimal known efficiency as well as for noise reduction. '''Rotor/stator gaps.''' Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. Design of stages is linked to the energy the blades have to give (compressor and fan) or take (turbine) to the air flow. To better define and understand that energy, we will use standard [http://en.wikipedia.org/wiki/State_function thermodynamic parameters] of gas, a.k.a state variables of a gas, on which are based quantities like the enthalpy: temperature, volume, pressure. We will also use the velocity because the actual work given by a turbofan engine is related to the mass flow rate of the gas expelled by the engine, which relates to velocity of this gas and the state variables. ===Gas variables: temperature, pressure, velocity=== ===Fan design=== Main goal of the fan is to increase the mass flow rate, mainly by increasing the velocity. The mass flow is linked to the area of the fan blades and the angular speed of the fan. To increase the velocity, simplest way is to reduce volume. The fan duct will thus have to act as a compressor on the aft-end. On the front-end, it is generally designed as an expander, to increase the pressure, allowing more efficient work on the air flow. ===Compressor design=== Main goal of the compressor is to increase pressure. Due to friction of the gas on the blades and guide vanes mainly, temperature is increased too. Thus, volume has to decrease, intake area will be greater than compressor discharge area. ===Turbine design=== Main goal of the turbine is to extract energy from the hot and fast gas discharged by the combustion into mechanical (rotational) work. Pressure and temperature may remain constant through the turbine, and high velocity and pressure will provide better efficiency ''[to be verified]''. ==Mechanical constraints== Blades on all three parts of a turbofan engine undergo heavy mechanical constraints due to high rpm achieved by the rotor, the high temperature in the turbine section and non-negligible temperature in the end of the compressor section, and the high pressure of the gas on which work is performed. Fan blades are not made of plain metal in real engines. In the eighties, they were made in honeycomb composite sandwich material, they are now made in triangular sandwich structure. For an engine of the size we are targeting, fan blades may be built with plain carbon fiber, like the [http://www.schuebeler-jets.com/index.php?option=com_content&task=view&id=102&Itemid=171 Schuebeler HST] high quality R/C engine. Compressor blades are made of titanium alloys, providing high strength and rigidity at these temperatures. Turbine blades are made of nickel alloys, better sustaining the high temperature, and still at higher strength than steel. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing] is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing hot isostatic pressing] possibly too, as explained on the ''How are made turbine blades'' [http://www.youtube.com/watch?v=vN3_Wkyl5PQ video]. I believe that a hot forging press can be done cheaply considering the small size of our blades. For the main fan, it thus may not be used. [[File:Blade.jpg|600px|center]] ''The above picture lacks the airfoil section of the blade, because the CAD software we use doesn't currently support it.'' Work-hardening and head-treating have to be studied, and depend on the crafting method used in the first place. The high-pressure turbine blades have to face very high temperature and pressure. On real engines, they are made of titanium and nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. Since the required lifetime is lower in our case, we may achieve a working engine with cheaper metals, like steel or nickel-rich alloy for the turbine blades. For the compressor blades, aluminum alloys are probably be a good solution. Don't forget that the blade insert will have to be milled at some point. [[Category:Turbofan|Blades]] 8b456dbc0c83e5810f1226b5ed08fb42fbf378f5 200 199 2011-05-09T22:00:23Z Vincent 1 /* Blade designs for efficient air flow */ fixing Griffith link and compressor parameters wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. ==Blade fixation== The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. In summary, real engines use a fir tree shape to hold the blade centrifugally, while keeping them free axially. They are generally fixed using a locking screw. This design allows for easy replacement of damaged blades, but is quite complex to manufacture, and requires lots of parts. Since our engine will be smaller, using fixing screws may not be feasible, and blades are subject to less centrifugal stress. Simpler blade fixation mechanism should be relevant. ==Blade designs for efficient air flow== Blades have to be carefully designed, because the overall efficiency of the engine largely depends on their design. '''Stages.''' They are not only made of blades on the rotor, but also blades on the stator, generally called ''vanes''. A '''stage''' is then a pair of a rotor stage and a stator stage, in this order. The stator prevents a rotating air flow to form inside the engine (swirl), driven by the action of the rotor blades. Stator vanes redirect the airflow in the more appropriate direction for the next rotor stage. They increase the energy of the gas ([http://en.wikipedia.org/wiki/Enthalpy enthalpy]) by removing the swirling effect that impairs it. '''Blade shape.''' Most basic design of a fan has flat-shaped blades. Twisted blades with a flat section are an improvement, taking into account the difference in apparent airflow velocity and torque all along the blade. Next step is to have a non-flat section, but an airfoil section. This has been [http://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines proved] to be the only design meeting efficiency requirements of the turbine engines, in 1926 by [http://en.wikipedia.org/wiki/Axial_compressor#Development Alan A. Griffith]. Finally, modern engines are designed with curved edges for the fan, for optimal known efficiency as well as for noise reduction. '''Rotor/stator gaps.''' Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. Design of stages is linked to the energy the blades have to give (compressor and fan) or take (turbine) to the air flow. To better define and understand that energy, we will use standard [http://en.wikipedia.org/wiki/State_function thermodynamic parameters] of gas, a.k.a state variables of a gas, on which are based quantities like the enthalpy: temperature, volume, pressure. We will also use the velocity because the actual work given by a turbofan engine is related to the mass flow rate of the gas expelled by the engine, which relates to velocity of this gas and the state variables. ===Gas variables: temperature, pressure, velocity=== ===Fan design=== The main goal of the fan is to increase the mass flow rate, mainly by increasing the velocity. The mass flow is related to the area of the fan blades and the angular speed of the fan. To increase the velocity, the simplest way is to reduce volume, in other words create a nozzle. The fan duct will thus have to act as a compressor on the aft-end. On the front-end, it is generally designed as an expander, to increase the pressure, allowing more efficient work on the air flow. ===Compressor design=== The main goal of the compressor is to increase pressure. A compressor stage is composed of a blade-mounted rotor and a vane-mounted stator. The shape of the blade is an airfoil and as the gas flows through the rotor and the stator it gains static pressure, since the blades form a expander. Speed is however gained in the rotor section, because of the high rotational speed of the blades. Stator vanes remove the resulting swirl, and converts the velocity (dynamic pressure) to static pressure, thus increasing again the pressure. The volume occupied by the gas can consequently drop as the pressure increases, intake area will be greater than compressor discharge area. Due to friction and pressure rise, temperature is increased too. ===Turbine design=== The main goal of the turbine is to extract energy from the hot and fast gas discharged by the combustion into mechanical (rotational) work. Pressure and temperature may remain constant through the turbine, and high velocity and pressure will provide better efficiency ''[to be verified]''. ==Mechanical constraints== Blades on all three parts of a turbofan engine undergo heavy mechanical constraints due to high rpm achieved by the rotor, the high temperature in the turbine section and non-negligible temperature in the end of the compressor section, and the high pressure of the gas on which work is performed. Fan blades are not made of plain metal in real engines. In the eighties, they were made in honeycomb composite sandwich material, they are now made in triangular sandwich structure. For an engine of the size we are targeting, fan blades may be built with plain carbon fiber, like the [http://www.schuebeler-jets.com/index.php?option=com_content&task=view&id=102&Itemid=171 Schuebeler HST] high quality R/C engine. Compressor blades are made of titanium alloys, providing high strength and rigidity at these temperatures. Turbine blades are made of nickel alloys, better sustaining the high temperature, and still at higher strength than steel. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing] is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing hot isostatic pressing] possibly too, as explained on the ''How are made turbine blades'' [http://www.youtube.com/watch?v=vN3_Wkyl5PQ video]. I believe that a hot forging press can be done cheaply considering the small size of our blades. For the main fan, it thus may not be used. [[File:Blade.jpg|600px|center]] ''The above picture lacks the airfoil section of the blade, because the CAD software we use doesn't currently support it.'' Work-hardening and head-treating have to be studied, and depend on the crafting method used in the first place. The high-pressure turbine blades have to face very high temperature and pressure. On real engines, they are made of titanium and nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. Since the required lifetime is lower in our case, we may achieve a working engine with cheaper metals, like steel or nickel-rich alloy for the turbine blades. For the compressor blades, aluminum alloys are probably be a good solution. Don't forget that the blade insert will have to be milled at some point. [[Category:Turbofan|Blades]] 55a6f849c9211ec75e72429ad26b9a81de6243af 0 11 152 147 2011-02-23T00:59:30Z Vincent 1 /* Our Design propositions */ wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, and it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. The issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Design considerations== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C... ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades, and maybe a simple design similar to the one above, but based on only one squared holder is enough. A completely different design of the mounting would be to use spacer parts between each blade, that would create the whole compressor stage fixation when all blades and spacer are put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more eavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages. Indeed, higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft. Creating a part that can be duplicated to form a cylinder, but with a non-parallel to the axis separation poses design and crafting issues. An alternative design would be to have the blades orthogonally free (axially too) and fix them on the sides using a key of some sort. The blades mounting ring and the blades would have a notch in the same place, in which a key would place to hold everything together. ===Fixing blades to stator=== To be studied. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ==Our Design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created to explain each subsystem and parts manufacturability. * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. [[Category:Turbofan]] 1ce46d20803da4accf3855f94c1a710c6c84f569 153 152 2011-02-23T01:03:06Z Vincent 1 category link down wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, and it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. The issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Design considerations== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C... ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades, and maybe a simple design similar to the one above, but based on only one squared holder is enough. A completely different design of the mounting would be to use spacer parts between each blade, that would create the whole compressor stage fixation when all blades and spacer are put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more eavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages. Indeed, higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft. Creating a part that can be duplicated to form a cylinder, but with a non-parallel to the axis separation poses design and crafting issues. An alternative design would be to have the blades orthogonally free (axially too) and fix them on the sides using a key of some sort. The blades mounting ring and the blades would have a notch in the same place, in which a key would place to hold everything together. ===Fixing blades to stator=== To be studied. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ==Our Design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. [[Category:Turbofan]] e31a4a48e0a4155b26f9efe9e5d2a84798296c00 155 153 2011-02-23T01:30:10Z Vincent 1 /* Design considerations */ wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, and it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. The issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Design considerations== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C. Cooling may be done by injecting low temperature air in the hot flow, or use film cooling in the combustion chamber. ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Providing power to the aircraft=== APUs and turbine engines provide power to aircrafts, either in an electric, hydraulic or pneumatic form. It would be nice to have an electrical power generator in our turbofan engines, because batteries are heavy. This is generally provided by the same mechanism than startup, used reversely, like an electric engine/alternator. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades, and maybe a simple design similar to the one above, but based on only one squared holder is enough. A completely different design of the mounting would be to use spacer parts between each blade, that would create the whole compressor stage fixation when all blades and spacer are put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages. Indeed, higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft. Creating a part that can be duplicated to form a cylinder, but with a non-parallel to the axis separation poses design and crafting issues. An alternative design would be to have the blades orthogonally free (axially too) and fix them on the sides using a key of some sort. The blades mounting ring and the blades would have a notch in the same place, in which a key would place to hold everything together. ===Fixing blades to stator=== To be studied. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ===Stator/rotor bearing=== Two kinds of bearings are used in turbines. * Stator and rotor are joint using a ball bearing constantly bathed in oil. * An oil bearing is used, in which pressurized oil prevents parts from touching. Carbon lip seals prevent the oil from escaping to the engine. ==Our Design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. [[Category:Turbofan]] e1f06f7ad8dc1e5ca3223139ebe1dced62cb7347 156 155 2011-02-23T01:36:03Z Vincent 1 /* Stator/rotor bearing */ wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, and it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. The issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Design considerations== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C. Cooling may be done by injecting low temperature air in the hot flow, or use film cooling in the combustion chamber. ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Providing power to the aircraft=== APUs and turbine engines provide power to aircrafts, either in an electric, hydraulic or pneumatic form. It would be nice to have an electrical power generator in our turbofan engines, because batteries are heavy. This is generally provided by the same mechanism than startup, used reversely, like an electric engine/alternator. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades, and maybe a simple design similar to the one above, but based on only one squared holder is enough. A completely different design of the mounting would be to use spacer parts between each blade, that would create the whole compressor stage fixation when all blades and spacer are put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages. Indeed, higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft. Creating a part that can be duplicated to form a cylinder, but with a non-parallel to the axis separation poses design and crafting issues. An alternative design would be to have the blades orthogonally free (axially too) and fix them on the sides using a key of some sort. The blades mounting ring and the blades would have a notch in the same place, in which a key would place to hold everything together. ===Fixing blades to stator=== To be studied. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ===Stator/rotor bearing=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance. Carbon lip seals prevent the oil from escaping to other parts of the engine. ==Our Design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. [[Category:Turbofan]] f83c30787885c225bec06107ec62372d409a074d 163 156 2011-02-24T00:03:11Z Vincent 1 /* Fixing blades to rotor */ moving to compressor's page wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofan advantages over other propelling ways are that they provide enough trust to climb at high altitude, possibly at attractive speed, and that they are not much fuel-greedy and thus can be running for some time with limited extra-weight. However they has the big disadvantage of being very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced, which will be the key for the decision to use turbofans or not. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage. The fan provides thrust from creating air pressure, and the combustion creates thrust by evacuating hot gas. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: it is often seen that a second concentric shaft for high pressure operations drives the low pressure shaft on which is mounted the fan. One stage engines exist and are less complicated and expensive to build, but are also less efficient. A gearbox may be needed to drive the fan if the low pressure shaft is still to fast. * The compression ratio is determined by the number of stages in the compressor and its efficiency. More compression mean more air to blend with fuel, and even more pressure at output, increasing the speed and mass of output gas, and thus overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A good book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, and it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. The issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Design considerations== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C. Cooling may be done by injecting low temperature air in the hot flow, or use film cooling in the combustion chamber. ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Providing power to the aircraft=== APUs and turbine engines provide power to aircrafts, either in an electric, hydraulic or pneumatic form. It would be nice to have an electrical power generator in our turbofan engines, because batteries are heavy. This is generally provided by the same mechanism than startup, used reversely, like an electric engine/alternator. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades, and maybe a simple design similar to the one above, but based on only one squared holder is enough. ===Fixing blades to stator=== To be studied. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ===Stator/rotor bearing=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance. Carbon lip seals prevent the oil from escaping to other parts of the engine. ==Our Design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. [[Category:Turbofan]] a167fe045bd9ce794456632729a68619cdfbc3e4 182 163 2011-05-02T18:45:02Z Vincent 1 /* How to build a cheap (~ $150) turbofan? */ rewriting parts wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets. They are however very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced while still having a reasonable efficiency, which is our primary concern here. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage feeding the combustion. The fan provides thrust from creating a massive air flow, and the turbine creates thrust by evacuating a hotter but less important air flow. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost has the fan mounted on. One stage engines exist and are less complicated and expensive to build, but are also less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas, and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but not a constraint as important as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing would be to have a simpler design of parts creating the internal volume of the turbine. In the above schema, we see that the shaft is straight and that the core envelope is curved to reduce volume on the high compression stage. If we take the same volumes on each part of the engine, and that we fix the envelope shape to a cylinder, the shaft has a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). A curved envelope is complicated to build, requiring lot of welding, but is used in real-world engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, and it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. The issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated part to build in a turbofan or turbojet engine is the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, driven by the action of compressor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. '''Open issues''': what is the most simple yet efficient shape for turbine and compressor fan blades? Is a flat shape acceptable? Blades need to overlap or not? Should they be build in a single piece of metal along with the axis mount ring or assembled from blades on mounting rings? '''How to manufacture the blades?''' ==Design considerations== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C. Cooling may be done by injecting low temperature air in the hot flow, or use film cooling in the combustion chamber. ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Providing power to the aircraft=== APUs and turbine engines provide power to aircrafts, either in an electric, hydraulic or pneumatic form. It would be nice to have an electrical power generator in our turbofan engines, because batteries are heavy. This is generally provided by the same mechanism than startup, used reversely, like an electric engine/alternator. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades, and maybe a simple design similar to the one above, but based on only one squared holder is enough. ===Fixing blades to stator=== To be studied. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ===Stator/rotor bearing=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance. Carbon lip seals prevent the oil from escaping to other parts of the engine. ==Our Design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. [[Category:Turbofan]] dc448228beef8b6d3d8d69883dcb3d624bf29293 183 182 2011-05-02T20:38:05Z Vincent 1 /* Design versus manufacturing */ blades... wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets. They are however very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced while still having a reasonable efficiency, which is our primary concern here. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage feeding the combustion. The fan provides thrust from creating a massive air flow, and the turbine creates thrust by evacuating a hotter but less important air flow. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost has the fan mounted on. One stage engines exist and are less complicated and expensive to build, but are also less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas, and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down because of the price. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing could be to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. Moreover, we may think about a turbine-stage mechanism embedded in the stator to try to cool it, which would make it hollow. The main issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine, because the gas generator (combustion) and the compressor can provide more power to the turbine. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, which would decrease the enthalpy of the gas (its internal energy), driven by the action of rotor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between rotor blades and the stator, as well as between the stator blades and the rotor. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. Blade geometric design by itself can reveal complicated. The first engine(s) had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1922, XXX proved that it blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to be built. Airfoils for blade design allow the compressor stages to better increase the velocity, since they provide a reducing area for the air to pass through (= a compressor), converter to pressure by stator blades. For turbine blades, it's the opposite, they provide a gas expander by increasing the area through which hot gases flow. ==Design considerations== ===Temperature control=== Cooling might be needed if low cost metals are used. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C. Cooling may be done by injecting low temperature air in the hot flow, or use film cooling in the combustion chamber. ===Startup=== Startup can be done at ground manually (with compressed air for example). Igniter has to be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten or something similar. The combustion should be self-igniting and self-maintaining, but if pumps or throttling lead to a discontinuous flow of fuel, the igniter will have to be available during the flight. ===Providing power to the aircraft=== APUs and turbine engines provide power to aircrafts, either in an electric, hydraulic or pneumatic form. It would be nice to have an electrical power generator in our turbofan engines, because batteries are heavy. This is generally provided by the same mechanism than startup, used reversely, like an electric engine/alternator. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades, and maybe a simple design similar to the one above, but based on only one squared holder is enough. ===Fixing blades to stator=== To be studied. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ===Stator/rotor bearing=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance. Carbon lip seals prevent the oil from escaping to other parts of the engine. ==Our Design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. [[Category:Turbofan]] 0580ef72bc324e80e263fc1a581d79daac15653a 184 183 2011-05-02T20:55:08Z Vincent 1 /* Design considerations */ temperature and ignition wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets. They are however very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced while still having a reasonable efficiency, which is our primary concern here. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage feeding the combustion. The fan provides thrust from creating a massive air flow, and the turbine creates thrust by evacuating a hotter but less important air flow. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost has the fan mounted on. One stage engines exist and are less complicated and expensive to build, but are also less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas, and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down because of the price. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing could be to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. Moreover, we may think about a turbine-stage mechanism embedded in the stator to try to cool it, which would make it hollow. The main issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine, because the gas generator (combustion) and the compressor can provide more power to the turbine. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, which would decrease the enthalpy of the gas (its internal energy), driven by the action of rotor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between rotor blades and the stator, as well as between the stator blades and the rotor. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. Blade geometric design by itself can reveal complicated. The first engine(s) had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1922, XXX proved that it blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to be built. Airfoils for blade design allow the compressor stages to better increase the velocity, since they provide a reducing area for the air to pass through (= a compressor), converter to pressure by stator blades. For turbine blades, it's the opposite, they provide a gas expander by increasing the area through which hot gases flow. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C, but will deform before melting. Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combusion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the stator of the turbine, which is, on actual engines, made of a ceramic coated high temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge circulates, enters the blades, and evacuates through drilled holes in the blades (convective cooling and film cooling again). For the rotor blades, the same principle is used, but with air coming from inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air, ignition would have to be made during the flight. ===Providing power to the aircraft=== APUs and turbine engines provide power to aircrafts, either in an electric, hydraulic or pneumatic form. It would be nice to have an electrical power generator in our turbofan engines, because batteries are heavy. This is generally provided by the same mechanism than startup, used reversely, like an electric engine/alternator. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades, and maybe a simple design similar to the one above, but based on only one squared holder is enough. ===Fixing blades to stator=== To be studied. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ===Stator/rotor bearing=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance. Carbon lip seals prevent the oil from escaping to other parts of the engine. ==Our Design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. [[Category:Turbofan]] 593f1c6b56f83c7dbc5e624c7babcd502e80bf71 185 184 2011-05-02T20:55:37Z Vincent 1 /* Temperature control */ typo wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets. They are however very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced while still having a reasonable efficiency, which is our primary concern here. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage feeding the combustion. The fan provides thrust from creating a massive air flow, and the turbine creates thrust by evacuating a hotter but less important air flow. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost has the fan mounted on. One stage engines exist and are less complicated and expensive to build, but are also less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas, and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down because of the price. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing could be to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. Moreover, we may think about a turbine-stage mechanism embedded in the stator to try to cool it, which would make it hollow. The main issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine, because the gas generator (combustion) and the compressor can provide more power to the turbine. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, which would decrease the enthalpy of the gas (its internal energy), driven by the action of rotor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between rotor blades and the stator, as well as between the stator blades and the rotor. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. Blade geometric design by itself can reveal complicated. The first engine(s) had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1922, XXX proved that it blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to be built. Airfoils for blade design allow the compressor stages to better increase the velocity, since they provide a reducing area for the air to pass through (= a compressor), converter to pressure by stator blades. For turbine blades, it's the opposite, they provide a gas expander by increasing the area through which hot gases flow. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C, but will deform before melting. Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the stator of the turbine, which is, on actual engines, made of a ceramic coated high temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge circulates, enters the blades, and evacuates through drilled holes in the blades (convective cooling and film cooling again). For the rotor blades, the same principle is used, but with air coming from inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air, ignition would have to be made during the flight. ===Providing power to the aircraft=== APUs and turbine engines provide power to aircrafts, either in an electric, hydraulic or pneumatic form. It would be nice to have an electrical power generator in our turbofan engines, because batteries are heavy. This is generally provided by the same mechanism than startup, used reversely, like an electric engine/alternator. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure. That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor, possibly using a magnet (not recommended due to the manufacturing process and temperatures it may face). Engine temperature should be recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade inserts gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades, and maybe a simple design similar to the one above, but based on only one squared holder is enough. ===Fixing blades to stator=== To be studied. ===External hardware=== Fuel tanks in the wings, fuel pumps, fuel lines, and engine mounting will have to be considered if turbofans are used. Sensors will require input ports on the computer, and pump driving (= engine control) will require at least one output port for each engine on the computer. ===Stator/rotor bearing=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance. Carbon lip seals prevent the oil from escaping to other parts of the engine. ==Our Design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. [[Category:Turbofan]] df984558904c1c05953c15e326c3fa46815aa4bc 186 185 2011-05-02T21:36:26Z Vincent 1 /* Design considerations */ the end of it wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets. They are however very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced while still having a reasonable efficiency, which is our primary concern here. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage feeding the combustion. The fan provides thrust from creating a massive air flow, and the turbine creates thrust by evacuating a hotter but less important air flow. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost has the fan mounted on. One stage engines exist and are less complicated and expensive to build, but are also less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas, and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down because of the price. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing could be to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. Moreover, we may think about a turbine-stage mechanism embedded in the stator to try to cool it, which would make it hollow. The main issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine, because the gas generator (combustion) and the compressor can provide more power to the turbine. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, which would decrease the enthalpy of the gas (its internal energy), driven by the action of rotor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between rotor blades and the stator, as well as between the stator blades and the rotor. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. Blade geometric design by itself can reveal complicated. The first engine(s) had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1922, XXX proved that it blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to be built. Airfoils for blade design allow the compressor stages to better increase the velocity, since they provide a reducing area for the air to pass through (= a compressor), converter to pressure by stator blades. For turbine blades, it's the opposite, they provide a gas expander by increasing the area through which hot gases flow. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C, but will deform before melting. Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the stator of the turbine, which is, on actual engines, made of a ceramic coated high temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge circulates, enters the blades, and evacuates through drilled holes in the blades (convective cooling and film cooling again). For the rotor blades, the same principle is used, but with air coming from inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air, ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. The rotor speed information would be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, maybe a simple square-section root is enough. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blades-roots are inserted and fixed. Creating a perpendicular root is already a challenge. Rotor's root would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting is '''to be considered'''. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Compressed air (a.k.a. bleed air) from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our Design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. [[Category:Turbofan]] c2abb3a86741948795334d1d65052a70c821c8af 192 186 2011-05-02T22:00:20Z Vincent 1 /* Our Design propositions */ adding combustor link wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets. They are however very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced while still having a reasonable efficiency, which is our primary concern here. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage feeding the combustion. The fan provides thrust from creating a massive air flow, and the turbine creates thrust by evacuating a hotter but less important air flow. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost has the fan mounted on. One stage engines exist and are less complicated and expensive to build, but are also less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas, and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down because of the price. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing could be to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. Moreover, we may think about a turbine-stage mechanism embedded in the stator to try to cool it, which would make it hollow. The main issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine, because the gas generator (combustion) and the compressor can provide more power to the turbine. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, which would decrease the enthalpy of the gas (its internal energy), driven by the action of rotor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between rotor blades and the stator, as well as between the stator blades and the rotor. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. Blade geometric design by itself can reveal complicated. The first engine(s) had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1922, XXX proved that it blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to be built. Airfoils for blade design allow the compressor stages to better increase the velocity, since they provide a reducing area for the air to pass through (= a compressor), converter to pressure by stator blades. For turbine blades, it's the opposite, they provide a gas expander by increasing the area through which hot gases flow. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C, but will deform before melting. Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the stator of the turbine, which is, on actual engines, made of a ceramic coated high temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge circulates, enters the blades, and evacuates through drilled holes in the blades (convective cooling and film cooling again). For the rotor blades, the same principle is used, but with air coming from inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air, ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. The rotor speed information would be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, maybe a simple square-section root is enough. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blades-roots are inserted and fixed. Creating a perpendicular root is already a challenge. Rotor's root would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting is '''to be considered'''. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Compressed air (a.k.a. bleed air) from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our Design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustor]]: Combustor are the power input of the engine, and need not to melt while maintaining the combustion. [[Category:Turbofan]] e438e571e0f093476f194342d24b2a1adbb0fac7 193 192 2011-05-02T22:00:47Z Vincent 1 /* Our Design propositions */ typo wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets. They are however very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced while still having a reasonable efficiency, which is our primary concern here. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage feeding the combustion. The fan provides thrust from creating a massive air flow, and the turbine creates thrust by evacuating a hotter but less important air flow. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost has the fan mounted on. One stage engines exist and are less complicated and expensive to build, but are also less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas, and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down because of the price. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing could be to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. Moreover, we may think about a turbine-stage mechanism embedded in the stator to try to cool it, which would make it hollow. The main issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine, because the gas generator (combustion) and the compressor can provide more power to the turbine. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, which would decrease the enthalpy of the gas (its internal energy), driven by the action of rotor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between rotor blades and the stator, as well as between the stator blades and the rotor. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. Blade geometric design by itself can reveal complicated. The first engine(s) had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1922, XXX proved that it blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to be built. Airfoils for blade design allow the compressor stages to better increase the velocity, since they provide a reducing area for the air to pass through (= a compressor), converter to pressure by stator blades. For turbine blades, it's the opposite, they provide a gas expander by increasing the area through which hot gases flow. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C, but will deform before melting. Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the stator of the turbine, which is, on actual engines, made of a ceramic coated high temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge circulates, enters the blades, and evacuates through drilled holes in the blades (convective cooling and film cooling again). For the rotor blades, the same principle is used, but with air coming from inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air, ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. The rotor speed information would be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, maybe a simple square-section root is enough. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blades-roots are inserted and fixed. Creating a perpendicular root is already a challenge. Rotor's root would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting is '''to be considered'''. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Compressed air (a.k.a. bleed air) from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our Design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustor are the power input of the engine, and need not to melt while maintaining the combustion. [[Category:Turbofan]] 91dbbe1a128cd0c959d93f0e0b48db9550e21cf3 194 193 2011-05-02T22:01:08Z Vincent 1 /* Our Design propositions */ typo wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets. They are however very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced while still having a reasonable efficiency, which is our primary concern here. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage feeding the combustion. The fan provides thrust from creating a massive air flow, and the turbine creates thrust by evacuating a hotter but less important air flow. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost has the fan mounted on. One stage engines exist and are less complicated and expensive to build, but are also less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas, and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down because of the price. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing could be to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. Moreover, we may think about a turbine-stage mechanism embedded in the stator to try to cool it, which would make it hollow. The main issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine, because the gas generator (combustion) and the compressor can provide more power to the turbine. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, which would decrease the enthalpy of the gas (its internal energy), driven by the action of rotor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between rotor blades and the stator, as well as between the stator blades and the rotor. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. Blade geometric design by itself can reveal complicated. The first engine(s) had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1922, XXX proved that it blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to be built. Airfoils for blade design allow the compressor stages to better increase the velocity, since they provide a reducing area for the air to pass through (= a compressor), converter to pressure by stator blades. For turbine blades, it's the opposite, they provide a gas expander by increasing the area through which hot gases flow. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C, but will deform before melting. Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the stator of the turbine, which is, on actual engines, made of a ceramic coated high temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge circulates, enters the blades, and evacuates through drilled holes in the blades (convective cooling and film cooling again). For the rotor blades, the same principle is used, but with air coming from inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air, ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. The rotor speed information would be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, maybe a simple square-section root is enough. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blades-roots are inserted and fixed. Creating a perpendicular root is already a challenge. Rotor's root would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting is '''to be considered'''. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Compressed air (a.k.a. bleed air) from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our Design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while maintaining the combustion. [[Category:Turbofan]] 252d8018f434d5c7bf5e91a415d39a529a86ec6d 196 194 2011-05-02T22:05:58Z Vincent 1 /* Our Design propositions */ typo wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets. They are however very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced while still having a reasonable efficiency, which is our primary concern here. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage feeding the combustion. The fan provides thrust from creating a massive air flow, and the turbine creates thrust by evacuating a hotter but less important air flow. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost has the fan mounted on. One stage engines exist and are less complicated and expensive to build, but are also less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas, and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down because of the price. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing could be to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. Moreover, we may think about a turbine-stage mechanism embedded in the stator to try to cool it, which would make it hollow. The main issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine, because the gas generator (combustion) and the compressor can provide more power to the turbine. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, which would decrease the enthalpy of the gas (its internal energy), driven by the action of rotor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between rotor blades and the stator, as well as between the stator blades and the rotor. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. Blade geometric design by itself can reveal complicated. The first engine(s) had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1922, XXX proved that it blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to be built. Airfoils for blade design allow the compressor stages to better increase the velocity, since they provide a reducing area for the air to pass through (= a compressor), converter to pressure by stator blades. For turbine blades, it's the opposite, they provide a gas expander by increasing the area through which hot gases flow. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C, but will deform before melting. Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the stator of the turbine, which is, on actual engines, made of a ceramic coated high temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge circulates, enters the blades, and evacuates through drilled holes in the blades (convective cooling and film cooling again). For the rotor blades, the same principle is used, but with air coming from inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air, ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. The rotor speed information would be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, maybe a simple square-section root is enough. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blades-roots are inserted and fixed. Creating a perpendicular root is already a challenge. Rotor's root would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting is '''to be considered'''. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Compressed air (a.k.a. bleed air) from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while maintaining the combustion. [[Category:Turbofan]] 9c79ccf5fab8c47f95fd592635d24874a71acc84 6 33 157 2011-02-23T22:50:25Z Vincent 1 Turbine or compressor blade, with its milled insert. wikitext text/x-wiki Turbine or compressor blade, with its milled insert. 01319b7aaee2cea22367f9f7b49ab76bc3e05a87 6 34 159 2011-02-23T23:05:54Z Vincent 1 Close-up view on the blade fixation system. wikitext text/x-wiki Close-up view on the blade fixation system. efe3877bcf19690ee746704de4eb751fa9b76ebe 6 35 160 2011-02-23T23:06:27Z Vincent 1 Close-up view on the blade fixation system (rear view). wikitext text/x-wiki Close-up view on the blade fixation system (rear view). 16fd23aef8cc99a10b24b4bc7f842a42fd566171 0 30 161 148 2011-02-23T23:22:20Z Vincent 1 /* Compressor blades mounting */ pictures wikitext text/x-wiki =Compressor design= Real-world engines can have nearly 20 compression stages. We will first build a 3-stage compressor, in a single-spool engine to assess how it can be scaled to a full power engine, or if it will require more stages (= quite costly). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. ==Rotor design== Special care has to be given to the rotor, since it will spin at very high rotation speeds (not calculated yet). A blade mounting failure will likely cause the loss of the aircraft, given the constraints. ===Compressor blades mounting=== For the first compressor stage and the fan, the design is simplified by putting the roots of the blades parallel to the shaft. On stages 2 and 3, blades are not long enough, and to finish at the correct angle and not be twisted too much, they need to start at a non-zero angle. That brings problems to the design of the blade socket and fixing. The two pictures below propose a solution to this problem: blades' roots remain linear, the milling of the mounting is linear too, while it's a round part. This makes a strange effect, but allows the blades to be properly fixed on their sides, and if not enough, by a keyed design. The pictures below represent this mounting system, in exploded view in the first row (larger versions available). [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the mounted blades and on the key system from the yellow part. [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ===2nd and 3rd stage's compressor blade close-up and manufacturing=== ''Coming soon.'' ==Stator design== Even more complicated. [[Category:Turbofan|Compressor]] 50fa347c769a7ed03b6b40e63ddc46a0d4d29ce4 162 161 2011-02-24T00:02:00Z Vincent 1 wikitext text/x-wiki =Compressor design= Real-world engines can have nearly 20 compression stages. We will first build a 3-stage compressor, in a single-spool engine to assess how it can be scaled to a full power engine, or if it will require more stages (= quite costly). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. ==Rotor design== Special care has to be given to the rotor, since it will spin at very high rotation speeds (not calculated yet). A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's fisrt stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This is illustrated in the pictures below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade socket and fixing: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more classical solution to this problem: the hub is a standalone part, in which are milled the inserts holes. Blades' inserts remain a linear part and the milling of the hub is thus linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly fixed on their sides, and if not enough, by a keyed design. The pictures below represent this mounting system, in exploded view in the first row (larger versions available). [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the mounted blades and on the key system from the yellow part. [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== Even more complicated. [[Category:Turbofan|Compressor]] 5fb8e4d595f10c8924a59a54b02b64489699cc91 164 162 2011-02-24T00:14:15Z Vincent 1 /* 2nd and 3rd stage's compressor blade close-up and manufacturing */ finished importing text from the base page. wikitext text/x-wiki =Compressor design= Real-world engines can have nearly 20 compression stages. We will first build a 3-stage compressor, in a single-spool engine to assess how it can be scaled to a full power engine, or if it will require more stages (= quite costly). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. ==Rotor design== Special care has to be given to the rotor, since it will spin at very high rotation speeds (not calculated yet). A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's fisrt stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This is illustrated in the pictures below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade insert and the hub: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more usual solution to this problem, based on a real standalone hub in which are milled the inserts holes. Blades' inserts would remain a linear part and the milling of the hub would thus be linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly and easily fixed. A keyed design could also be added on the extremities of the blades, i.e. the side of the hub. The pictures below represent this mounting system, in exploded view (larger versions available). The hub is the grey round part with only one insert milled in it, the yellow part is the inter-stage spacer that has the key holding the blades. [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the blades mounted on the hub (in orange) and on the key system from the inter-stage spacer (in yellow). [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== Even more complicated. [[Category:Turbofan|Compressor]] 618f9fe5e28c3c6e782bc81cfd5d1b2007135581 165 164 2011-02-24T00:16:19Z Vincent 1 /* Stator design */ wikitext text/x-wiki =Compressor design= Real-world engines can have nearly 20 compression stages. We will first build a 3-stage compressor, in a single-spool engine to assess how it can be scaled to a full power engine, or if it will require more stages (= quite costly). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. ==Rotor design== Special care has to be given to the rotor, since it will spin at very high rotation speeds (not calculated yet). A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's fisrt stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This is illustrated in the pictures below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade insert and the hub: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more usual solution to this problem, based on a real standalone hub in which are milled the inserts holes. Blades' inserts would remain a linear part and the milling of the hub would thus be linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly and easily fixed. A keyed design could also be added on the extremities of the blades, i.e. the side of the hub. The pictures below represent this mounting system, in exploded view (larger versions available). The hub is the grey round part with only one insert milled in it, the yellow part is the inter-stage spacer that has the key holding the blades. [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the blades mounted on the hub (in orange) and on the key system from the inter-stage spacer (in yellow). [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== The stator will have as main issue the fact that blades must have a very thin insert. Shaping them like a T should be considered. [[Category:Turbofan|Compressor]] 7c8c1377f6ea6d4bc083db54dd47121a8dff0dc2 166 165 2011-02-24T00:19:17Z Vincent 1 /* Fan and compressor's fisrt stage blades mounting */ wikitext text/x-wiki =Compressor design= Real-world engines can have nearly 20 compression stages. We will first build a 3-stage compressor, in a single-spool engine to assess how it can be scaled to a full power engine, or if it will require more stages (= quite costly). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. ==Rotor design== Special care has to be given to the rotor, since it will spin at very high rotation speeds (not calculated yet). A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's fisrt stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. This design is depicted below. '''''PICTURE NEEDED''''' An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This keyed design is also considered for higher stages, and is illustrated in the models below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade insert and the hub: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more usual solution to this problem, based on a real standalone hub in which are milled the inserts holes. Blades' inserts would remain a linear part and the milling of the hub would thus be linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly and easily fixed. A keyed design could also be added on the extremities of the blades, i.e. the side of the hub. The pictures below represent this mounting system, in exploded view (larger versions available). The hub is the grey round part with only one insert milled in it, the yellow part is the inter-stage spacer that has the key holding the blades. [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the blades mounted on the hub (in orange) and on the key system from the inter-stage spacer (in yellow). [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== The stator will have as main issue the fact that blades must have a very thin insert. Shaping them like a T should be considered. [[Category:Turbofan|Compressor]] 84654ed35a46ea11deacac46f5e84a1b6fdeb818 169 166 2011-02-28T19:55:22Z Vincent 1 Adding compressor pictures wikitext text/x-wiki =Compressor design= Real-world engines can have nearly 20 compression stages. We will first build a 3-stage compressor, in a single-spool engine to assess how it can be scaled to a full power engine, or if it will require more stages (= quite costly). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. The two pictures below show an overall sketching of the compressor with all its blades. [[File:Compressor_noshell.jpg|300px]] [[File:Compressor_side.jpg|300px]] ==Rotor design== Special care has to be given to the rotor, since it will spin at very high rotation speeds (not calculated yet). A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's fisrt stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. This design is depicted below. '''''PICTURE NEEDED''''' An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This keyed design is also considered for higher stages, and is illustrated in the models below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade insert and the hub: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more usual solution to this problem, based on a real standalone hub in which are milled the inserts holes. Blades' inserts would remain a linear part and the milling of the hub would thus be linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly and easily fixed. A keyed design could also be added on the extremities of the blades, i.e. the side of the hub. The pictures below represent this mounting system, in exploded view (larger versions available). The hub is the grey round part with only one insert milled in it, the yellow part is the inter-stage spacer that has the key holding the blades. [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the blades mounted on the hub (in orange) and on the key system from the inter-stage spacer (in yellow). [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== The stator will have as main issue the fact that blades must have a very thin insert. Shaping them like a T should be considered. [[Category:Turbofan|Compressor]] 51abdacb6498795178299f842c576e9148ae40be 174 169 2011-02-28T20:11:39Z Vincent 1 Rotor and stator pictures wikitext text/x-wiki =Compressor design= Real-world engines can have nearly 20 compression stages. We will first build a 3-stage compressor, in a single-spool engine to assess how it can be scaled to a full power engine, or if it will require more stages (= quite costly). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. The two pictures below show an overall sketching of the compressor with all its blades. [[File:Compressor_noshell.jpg|300px]] [[File:Compressor_side.jpg|300px]] The three stages depicted here have the following number of blades (same for rotor and stator of the same stage): * Stage1: 18 (20° spaced) * Stage2: 20 (18° spaced) * Stage3: 24 (15° spaced) ==Rotor design== The following pictures represent the shaft and the three stages compressor rotor blades. [[File:Rotor_front_side.jpg|300px]] [[File:Rotor_side.jpg|300px]] Special care has to be given to the rotor, since it will spin at very high rotation speeds (not calculated yet). A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's fisrt stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. This design is depicted below. '''''PICTURE NEEDED''''' An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This keyed design is also considered for higher stages, and is illustrated in the models below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade insert and the hub: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more usual solution to this problem, based on a real standalone hub in which are milled the inserts holes. Blades' inserts would remain a linear part and the milling of the hub would thus be linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly and easily fixed. A keyed design could also be added on the extremities of the blades, i.e. the side of the hub. The pictures below represent this mounting system, in exploded view (larger versions available). The hub is the grey round part with only one insert milled in it, the yellow part is the inter-stage spacer that has the key holding the blades. [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the blades mounted on the hub (in orange) and on the key system from the inter-stage spacer (in yellow). [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== The stator will have as main issue the fact that blades must have a very thin insert. Shaping them like a T should be considered. [[File:Stator_side.jpg|300px]] [[File:Stator_front.jpg|300px]] [[Category:Turbofan|Compressor]] 457a2b678ef4a5dd65c0ddb53b5aaafd77487bfd 179 174 2011-02-28T20:59:47Z Vincent 1 wikitext text/x-wiki =Compressor design= Real-world engines can have nearly 20 compression stages. We will first build a 3-stage compressor, in a single-spool engine to assess how it can be scaled to a full power engine, or if it will require more stages (= quite costly). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. The two pictures below show an overall sketching of the compressor with all its blades. [[File:Compressor_noshell.jpg|300px]] [[File:Compressor_side.jpg|300px]] The three stages depicted here have the following number of blades (same for rotor and stator of the same stage): * Stage1: 18 (20° spaced) * Stage2: 20 (18° spaced) * Stage3: 24 (15° spaced) ==Rotor design== The following pictures represent the shaft and the three stages compressor rotor blades. [[File:Rotor_front_side.jpg|300px]] [[File:Rotor_side.jpg|300px]] Special care has to be given to the rotor, since it will spin at very high rotation speeds (not calculated yet). A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's fisrt stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. This design is depicted below. [[File:Blade_fixation_simple.jpg|300px|center]] An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This keyed design is also considered for higher stages, and is illustrated in the models below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade insert and the hub: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more usual solution to this problem, based on a real standalone hub in which are milled the inserts holes. Blades' inserts would remain a linear part and the milling of the hub would thus be linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly and easily fixed. A keyed design could also be added on the extremities of the blades, i.e. the side of the hub. The pictures below represent this mounting system, in exploded view (larger versions available). The hub is the grey round part with only one insert milled in it, the yellow part is the inter-stage spacer that has the key holding the blades. [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the blades mounted on the hub (in orange) and on the key system from the inter-stage spacer (in yellow). [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== The stator will have as main issue the fact that blades must have a very thin insert. Shaping them like a T should be considered. [[File:Stator_side.jpg|300px]] [[File:Stator_front.jpg|300px]] [[Category:Turbofan|Compressor]] 9baf8a9c5b2155d4aa2688601628edd4d45beba6 197 179 2011-05-02T22:16:48Z Vincent 1 /* Compressor design */ typos and fixes wikitext text/x-wiki =Compressor design= Real-world engines have 10 to 20 compression stages. We will first build a 3- to 5-stage compressor, in a single-spool engine to assess how it can be scaled to a full power engine, or if it will require more stages (= more expensive). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. The two pictures below show an overall sketching of the compressor with all its blades, flat rendered, but will be updated to be airfoil rendered someday if we can find how to draw it. [[File:Compressor_noshell.jpg|300px]] [[File:Compressor_side.jpg|300px]] The three stages depicted here have the following number of blades (same for rotor and stator of the same stage): * Stage1: 18 (20° spaced) * Stage2: 20 (18° spaced) * Stage3: 24 (15° spaced) ==Rotor design== The following pictures represent the shaft and the three stages compressor rotor blades. [[File:Rotor_front_side.jpg|300px]] [[File:Rotor_side.jpg|300px]] Special care has to be given to the rotor, since it will spin at very high rotation speeds ('''to be calculated'''). A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's first stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. This design is depicted below. [[File:Blade_fixation_simple.jpg|300px|center]] An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This keyed design is also considered for higher stages, and is illustrated in the models below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade insert and the hub: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more usual solution to this problem, based on a real standalone hub in which are milled the inserts holes. Blades' inserts would remain a linear part and the milling of the hub would thus be linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly and easily fixed. A keyed design could also be added on the extremities of the blades, i.e. the side of the hub. The pictures below represent this mounting system, in exploded view (larger versions available). The hub is the grey round part with only one insert milled in it, the yellow part is the inter-stage spacer that has the key holding the blades. [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the blades mounted on the hub (in orange) and on the key system from the inter-stage spacer (in yellow). [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== The stator will have as main issue the fact that blades must have a very thin insert. Shaping them like a T should be considered, but bolting them on the stator looks more promising for now. [[File:Stator_side.jpg|300px]] [[File:Stator_front.jpg|300px]] [[Category:Turbofan|Compressor]] 429520c61ca47bdf5fa2128492586a87918f0c58 6 36 167 2011-02-28T19:46:06Z Vincent 1 The 3-stage compressor with all blades mounted, seen from the side and without its shell. wikitext text/x-wiki The 3-stage compressor with all blades mounted, seen from the side and without its shell. 165b86c7ff9524cd693568e82d5991f1a2a8aab2 176 167 2011-02-28T20:23:16Z Vincent 1 uploaded a new version of "[[File:Compressor side.jpg]]":&#32;yellow shaft wikitext text/x-wiki The 3-stage compressor with all blades mounted, seen from the side and without its shell. 165b86c7ff9524cd693568e82d5991f1a2a8aab2 6 37 168 2011-02-28T19:46:57Z Vincent 1 The 3-stage compressor with all blades mounted without its shell. wikitext text/x-wiki The 3-stage compressor with all blades mounted without its shell. 2ce5e15af437fb6bd0208a96e782ef19e27db574 175 168 2011-02-28T20:20:24Z Vincent 1 uploaded a new version of "[[File:Compressor noshell.jpg]]" wikitext text/x-wiki The 3-stage compressor with all blades mounted without its shell. 2ce5e15af437fb6bd0208a96e782ef19e27db574 6 38 170 2011-02-28T19:56:48Z Vincent 1 Stator blades mounted on their shell and the shaft seen from the front. wikitext text/x-wiki Stator blades mounted on their shell and the shaft seen from the front. e97b092ad9c13a25a61f73f6441e42d3132596d6 6 39 171 2011-02-28T19:57:50Z Vincent 1 Stator blades mounted on their shell and the shaft profile visible, nearly seen from the side. wikitext text/x-wiki Stator blades mounted on their shell and the shaft profile visible, nearly seen from the side. 35d2051d7612c39bf27eaeeeaceee8fbf12f1053 6 40 172 2011-02-28T20:08:34Z Vincent 1 Rotor with its three-stages blades, seen from nearly front. wikitext text/x-wiki Rotor with its three-stages blades, seen from nearly front. 93760bbb7cd309452b030ad9193ce705038c8899 6 41 173 2011-02-28T20:09:09Z Vincent 1 Rotor with its three-stages blades seen from the side. wikitext text/x-wiki Rotor with its three-stages blades seen from the side. d228ed9be4c155d2f97ba8baa35c490318df7397 6 42 177 2011-02-28T20:56:30Z Vincent 1 Simple way of fixing blades: insert is parallel to the shaft and a spacer is used to hold them together, actually composing the fan hub. wikitext text/x-wiki Simple way of fixing blades: insert is parallel to the shaft and a spacer is used to hold them together, actually composing the fan hub. 101a8d9483b795937c94901cea914384126790bd 178 177 2011-02-28T20:58:31Z Vincent 1 uploaded a new version of "[[File:Blade fixation simple.jpg]]":&#32;fixing display artefact wikitext text/x-wiki Simple way of fixing blades: insert is parallel to the shaft and a spacer is used to hold them together, actually composing the fan hub. 101a8d9483b795937c94901cea914384126790bd 0 16 180 128 2011-03-30T10:54:11Z Vincent 1 /* Rocket engines */ Huzel and Huang wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Web pages== * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Books (online links)== ===Multi-domain=== * [http://books.google.fr/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.fr/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.fr/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific parts:==== * [http://books.google.fr/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.fr/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.fr/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.fr/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.fr/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.fr/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.fr/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. 7f7b90b6b9083865d88d53db7048cdd72fd7b1e3 181 180 2011-03-31T01:01:40Z Vincent 1 /* Web pages */ arocketry links wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/l amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Books (online links)== ===Multi-domain=== * [http://books.google.fr/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.fr/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.fr/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific parts:==== * [http://books.google.fr/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.fr/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.fr/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.fr/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.fr/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.fr/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.fr/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. 752c206d4bb0663c30913553121ec65e7fccf130 0 43 187 2011-05-02T21:54:20Z Vincent 1 first release of the aft-hybrid fan wikitext text/x-wiki =Alternative design for turbofans= All recent engines have the same basic architecture: a fan at the front, the turbine engine below it, and the two flows mix at the exhaust. Early designs of turbofans were actually created by putting a ducted fan on the aft part of a turbojet engine, since turbojet engines were already existing. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a quite hot gas flow, which would eventually corrode or fatigue it more than at normal temperature. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or low power design include a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans, at least as long as research or people won't have proven it was wrong: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. The so-called hybrid turbine would be a mix of axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its course, while using (a part of) the energy required to divert it. The aft-fan would intake the mixed flow of the fresh intake and the turbine discharge, providing higher energy to the fan flow. Properly mixing the two flows would allow the fan to be build with low enough temperature metals, like aluminum alloys. The result will inevitably be a slightly longer engine. ''Drawings (schematics or 3D CAD models) are coming soon'' 328b2856e4f0217003fed68fe9e429e77510d744 188 187 2011-05-02T21:55:12Z Vincent 1 /* Alternative design for turbofans */ wikitext text/x-wiki =Alternative design for turbofans= All recent engines have the same basic architecture: a fan at the front, the turbine engine below it, and the two flows mix at the exhaust. Early designs of turbofans were actually created by putting a ducted fan on the aft part of a turbojet engine, since turbojet engines were already existing. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a quite hot gas flow, which would eventually corrode or fatigue it more than at normal temperature. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or low power design include a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans, at least as long as research or people won't have proven it was wrong: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. The so-called hybrid turbine would be a mix of axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its course, while using (a part of) the energy required to divert it. The aft-fan would intake the mixed flow of the fresh intake and the turbine discharge, providing higher energy to the fan flow. Properly mixing the two flows would allow the fan to be build with low enough temperature metals, like aluminum alloys. The result will inevitably be a slightly longer engine. ''Drawings (schematics or 3D CAD models) are coming soon'' [[Category:Turbofan|Compressor]] 65b5908bcdd1ea3f26fc34c02d62a7e28e01bf7d 189 188 2011-05-02T21:56:18Z Vincent 1 category wikitext text/x-wiki =Alternative design for turbofans= All recent engines have the same basic architecture: a fan at the front, the turbine engine below it, and the two flows mix at the exhaust. Early designs of turbofans were actually created by putting a ducted fan on the aft part of a turbojet engine, since turbojet engines were already existing. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a quite hot gas flow, which would eventually corrode or fatigue it more than at normal temperature. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or low power design include a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans, at least as long as research or people won't have proven it was wrong: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. The so-called hybrid turbine would be a mix of axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its course, while using (a part of) the energy required to divert it. The aft-fan would intake the mixed flow of the fresh intake and the turbine discharge, providing higher energy to the fan flow. Properly mixing the two flows would allow the fan to be build with low enough temperature metals, like aluminum alloys. The result will inevitably be a slightly longer engine. ''Drawings (schematics or 3D CAD models) are coming soon.'' [[Category:Turbofan|Innovative design]] 3006dbba9a12960b4ffd3ace7f746602f0cf651f 190 189 2011-05-02T21:56:59Z Vincent 1 wikitext text/x-wiki =Alternative design for turbofans= All recent engines have the same basic architecture: a fan at the front, the turbine engine below it, and the two flows mix at the exhaust. Early designs of turbofans were actually created by putting a ducted fan on the aft part of a turbojet engine, since turbojet engines were already existing. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a quite hot gas flow, which would eventually corrode or fatigue it more than at normal temperature. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or low power design include a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans, at least as long as research or people won't have proven it was wrong: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. The so-called hybrid turbine would be a mix of axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its course, while using (a part of) the energy required to divert it. The aft-fan would intake the mixed flow of the fresh intake and the turbine discharge, providing higher energy to the fan flow. Properly mixing the two flows would allow the fan to be build with low enough temperature metals, like aluminum alloys. The result will inevitably be a slightly longer engine. ''Drawings (schematics or 3D CAD models) are coming soon.'' [[Category:Turbofan]] ae0206662623c38d5151bb9e5e28e39ba384afdb 191 190 2011-05-02T21:57:22Z Vincent 1 wikitext text/x-wiki =Alternative design for turbofans= All recent engines have the same basic architecture: a fan at the front, the turbine engine below it, and the two flows mix at the exhaust. Early designs of turbofans were actually created by putting a ducted fan on the aft part of a turbojet engine, since turbojet engines were already existing. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a quite hot gas flow, which would eventually corrode or fatigue it more than at normal temperature. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or low power design include a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans, at least as long as research or people won't have proven it was wrong: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. The so-called hybrid turbine would be a mix of axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its course, while using (a part of) the energy required to divert it. The aft-fan would intake the mixed flow of the fresh intake and the turbine discharge, providing higher energy to the fan flow. Properly mixing the two flows would allow the fan to be build with low enough temperature metals, like aluminum alloys. The result will inevitably be a slightly longer engine. ''Drawings (schematics or 3D CAD models) are coming soon.'' [[Category:Turbofan|Alternative design]] 2cd8dfcb58f306ca043907c513b92e40cf77fe49 0 44 195 2011-05-02T22:04:04Z Vincent 1 draft for combustors wikitext text/x-wiki =Turbofan combustors= ==Typical designs== ===Can combustor=== ===Annular combustors=== ===Can-annular combustors=== ==Our design== ===Flame holding=== ===Cooling=== ===Material=== ===Ignition=== [[Category:Turbofan|Combustors]] ae503c120a240a1aedbe1a53a348202fd47707c9 6 45 201 2011-05-09T23:20:05Z Vincent 1 Compressor blades 2D schematics (source SVG file). wikitext text/x-wiki Compressor blades 2D schematics (source SVG file). 41751910e468e2ff930b7795187aefffcbb33a0e 205 201 2011-05-09T23:27:41Z Vincent 1 uploaded a new version of "[[File:Compressor blades.svg]]":&#32;white bg wikitext text/x-wiki Compressor blades 2D schematics (source SVG file). 41751910e468e2ff930b7795187aefffcbb33a0e 6 46 202 2011-05-09T23:20:45Z Vincent 1 Compressor blades 2D schematics. wikitext text/x-wiki Compressor blades 2D schematics. e4fc283dbbe55733303c01c6ba40834509539a99 204 202 2011-05-09T23:26:02Z Vincent 1 uploaded a new version of "[[File:Compressor blades.png]]":&#32;white background wikitext text/x-wiki Compressor blades 2D schematics. e4fc283dbbe55733303c01c6ba40834509539a99 0 32 203 200 2011-05-09T23:24:00Z Vincent 1 Adding link to compressor blades image. wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. ==Blade fixation== The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. In summary, real engines use a fir tree shape to hold the blade centrifugally, while keeping them free axially. They are generally fixed using a locking screw. This design allows for easy replacement of damaged blades, but is quite complex to manufacture, and requires lots of parts. Since our engine will be smaller, using fixing screws may not be feasible, and blades are subject to less centrifugal stress. Simpler blade fixation mechanism should be relevant. ==Blade designs for efficient air flow== Blades have to be carefully designed, because the overall efficiency of the engine largely depends on their design. '''Stages.''' They are not only made of blades on the rotor, but also blades on the stator, generally called ''vanes''. A '''stage''' is then a pair of a rotor stage and a stator stage, in this order. The stator prevents a rotating air flow to form inside the engine (swirl), driven by the action of the rotor blades. Stator vanes redirect the airflow in the more appropriate direction for the next rotor stage. They increase the energy of the gas ([http://en.wikipedia.org/wiki/Enthalpy enthalpy]) by removing the swirling effect that impairs it. '''Blade shape.''' Most basic design of a fan has flat-shaped blades. Twisted blades with a flat section are an improvement, taking into account the difference in apparent airflow velocity and torque all along the blade. Next step is to have a non-flat section, but an airfoil section. This has been [http://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines proved] to be the only design meeting efficiency requirements of the turbine engines, in 1926 by [http://en.wikipedia.org/wiki/Axial_compressor#Development Alan A. Griffith]. Finally, modern engines are designed with curved edges for the fan, for optimal known efficiency as well as for noise reduction. '''Rotor/stator gaps.''' Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. Design of stages is linked to the energy the blades have to give (compressor and fan) or take (turbine) to the air flow. To better define and understand that energy, we will use standard [http://en.wikipedia.org/wiki/State_function thermodynamic parameters] of gas, a.k.a state variables of a gas, on which are based quantities like the enthalpy: temperature, volume, pressure. We will also use the velocity because the actual work given by a turbofan engine is related to the mass flow rate of the gas expelled by the engine, which relates to velocity of this gas and the state variables. ===Gas variables: temperature, pressure, velocity=== ===Fan design=== The main goal of the fan is to increase the mass flow rate, mainly by increasing the velocity. The mass flow is related to the area of the fan blades and the angular speed of the fan. To increase the velocity, the simplest way is to reduce volume, in other words create a nozzle. The fan duct will thus have to act as a compressor on the aft-end. On the front-end, it is generally designed as an expander, to increase the pressure, allowing more efficient work on the air flow. ===Compressor design=== The main goal of the compressor is to increase pressure. A compressor stage is composed of a blade-mounted rotor and a vane-mounted stator. The shape of the blade is an airfoil and as the gas flows through the rotor and the stator it gains static pressure, since the blades form a expander. Speed is however gained in the rotor section, because of the high rotational speed of the blades. Stator vanes remove the resulting swirl, and converts the velocity (dynamic pressure) to static pressure, thus increasing again the pressure. The volume occupied by the gas can consequently drop as the pressure increases, intake area will be greater than compressor discharge area. Due to friction and pressure rise, temperature is increased too. [[File:Compressor_blades.png|360px|center]] ===Turbine design=== The main goal of the turbine is to extract energy from the hot and fast gas discharged by the combustion into mechanical (rotational) work. Pressure and temperature may remain constant through the turbine, and high velocity and pressure will provide better efficiency ''[to be verified]''. ==Mechanical constraints== Blades on all three parts of a turbofan engine undergo heavy mechanical constraints due to high rpm achieved by the rotor, the high temperature in the turbine section and non-negligible temperature in the end of the compressor section, and the high pressure of the gas on which work is performed. Fan blades are not made of plain metal in real engines. In the eighties, they were made in honeycomb composite sandwich material, they are now made in triangular sandwich structure. For an engine of the size we are targeting, fan blades may be built with plain carbon fiber, like the [http://www.schuebeler-jets.com/index.php?option=com_content&task=view&id=102&Itemid=171 Schuebeler HST] high quality R/C engine. Compressor blades are made of titanium alloys, providing high strength and rigidity at these temperatures. Turbine blades are made of nickel alloys, better sustaining the high temperature, and still at higher strength than steel. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing] is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing hot isostatic pressing] possibly too, as explained on the ''How are made turbine blades'' [http://www.youtube.com/watch?v=vN3_Wkyl5PQ video]. I believe that a hot forging press can be done cheaply considering the small size of our blades. For the main fan, it thus may not be used. [[File:Blade.jpg|600px|center]] ''The above picture lacks the airfoil section of the blade, because the CAD software we use doesn't currently support it.'' Work-hardening and head-treating have to be studied, and depend on the crafting method used in the first place. The high-pressure turbine blades have to face very high temperature and pressure. On real engines, they are made of titanium and nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. Since the required lifetime is lower in our case, we may achieve a working engine with cheaper metals, like steel or nickel-rich alloy for the turbine blades. For the compressor blades, aluminum alloys are probably be a good solution. Don't forget that the blade insert will have to be milled at some point. [[Category:Turbofan|Blades]] 564e50e473e6e7b800d31e0ec7b13f8df707ebdf 208 203 2011-07-27T23:20:42Z Vincent 1 /* Manufacturing propositions */ wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. ==Blade fixation== The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. In summary, real engines use a fir tree shape to hold the blade centrifugally, while keeping them free axially. They are generally fixed using a locking screw. This design allows for easy replacement of damaged blades, but is quite complex to manufacture, and requires lots of parts. Since our engine will be smaller, using fixing screws may not be feasible, and blades are subject to less centrifugal stress. Simpler blade fixation mechanism should be relevant. ==Blade designs for efficient air flow== Blades have to be carefully designed, because the overall efficiency of the engine largely depends on their design. '''Stages.''' They are not only made of blades on the rotor, but also blades on the stator, generally called ''vanes''. A '''stage''' is then a pair of a rotor stage and a stator stage, in this order. The stator prevents a rotating air flow to form inside the engine (swirl), driven by the action of the rotor blades. Stator vanes redirect the airflow in the more appropriate direction for the next rotor stage. They increase the energy of the gas ([http://en.wikipedia.org/wiki/Enthalpy enthalpy]) by removing the swirling effect that impairs it. '''Blade shape.''' Most basic design of a fan has flat-shaped blades. Twisted blades with a flat section are an improvement, taking into account the difference in apparent airflow velocity and torque all along the blade. Next step is to have a non-flat section, but an airfoil section. This has been [http://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines proved] to be the only design meeting efficiency requirements of the turbine engines, in 1926 by [http://en.wikipedia.org/wiki/Axial_compressor#Development Alan A. Griffith]. Finally, modern engines are designed with curved edges for the fan, for optimal known efficiency as well as for noise reduction. '''Rotor/stator gaps.''' Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by the engines. Design of stages is linked to the energy the blades have to give (compressor and fan) or take (turbine) to the air flow. To better define and understand that energy, we will use standard [http://en.wikipedia.org/wiki/State_function thermodynamic parameters] of gas, a.k.a state variables of a gas, on which are based quantities like the enthalpy: temperature, volume, pressure. We will also use the velocity because the actual work given by a turbofan engine is related to the mass flow rate of the gas expelled by the engine, which relates to velocity of this gas and the state variables. ===Gas variables: temperature, pressure, velocity=== ===Fan design=== The main goal of the fan is to increase the mass flow rate, mainly by increasing the velocity. The mass flow is related to the area of the fan blades and the angular speed of the fan. To increase the velocity, the simplest way is to reduce volume, in other words create a nozzle. The fan duct will thus have to act as a compressor on the aft-end. On the front-end, it is generally designed as an expander, to increase the pressure, allowing more efficient work on the air flow. ===Compressor design=== The main goal of the compressor is to increase pressure. A compressor stage is composed of a blade-mounted rotor and a vane-mounted stator. The shape of the blade is an airfoil and as the gas flows through the rotor and the stator it gains static pressure, since the blades form a expander. Speed is however gained in the rotor section, because of the high rotational speed of the blades. Stator vanes remove the resulting swirl, and converts the velocity (dynamic pressure) to static pressure, thus increasing again the pressure. The volume occupied by the gas can consequently drop as the pressure increases, intake area will be greater than compressor discharge area. Due to friction and pressure rise, temperature is increased too. [[File:Compressor_blades.png|360px|center]] ===Turbine design=== The main goal of the turbine is to extract energy from the hot and fast gas discharged by the combustion into mechanical (rotational) work. Pressure and temperature may remain constant through the turbine, and high velocity and pressure will provide better efficiency ''[to be verified]''. ==Mechanical constraints== Blades on all three parts of a turbofan engine undergo heavy mechanical constraints due to high rpm achieved by the rotor, the high temperature in the turbine section and non-negligible temperature in the end of the compressor section, and the high pressure of the gas on which work is performed. Fan blades are not made of plain metal in real engines. In the eighties, they were made in honeycomb composite sandwich material, they are now made in triangular sandwich structure. For an engine of the size we are targeting, fan blades may be built with plain carbon fiber, like the [http://www.schuebeler-jets.com/index.php?option=com_content&task=view&id=102&Itemid=171 Schuebeler HST] high quality R/C engine. Compressor blades are made of titanium alloys, providing high strength and rigidity at these temperatures. Turbine blades are made of nickel alloys, better sustaining the high temperature, and still at higher strength than steel. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing], as a [http://en.wikipedia.org/wiki/Hot_working Hot working] process, is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing hot isostatic pressing] possibly too, as explained on the ''How are made turbine blades'' [http://www.youtube.com/watch?v=vN3_Wkyl5PQ video]. I believe that a hot forging press can be done cheaply considering the small size of our blades. For the main fan, it thus may not be used. [[File:Blade.jpg|600px|center]] ''The above picture lacks the airfoil section of the blade, because the CAD software we use doesn't currently support it.'' Work-hardening and heat-resistance processing have to be studied, and depend on the crafting method used in the first place. The high-pressure turbine blades have to face very high temperature and pressure. On real engines, they are made of titanium and nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. Since the required lifetime is lower in our case, we may achieve a working engine with cheaper metals, like steel or nickel-rich alloys for the turbine blades. Historically, the temperature that turbine blades could sustain greatly improved over time, as well as engine efficiency. We can take these two factors and design a reasonably short-lived (6hrs, 3 missions?) and medium efficiency engine. Material for compressor blade and fan blade can probably be aluminum or aluminum alloys like 2024 (dural) or even 7075. The blade insert will have to be milled at some point, unless if it is casted or pressed. Sonic speed at the tip of a 160mm fan is achieved at around 40000 rpm, for 120mm a fan, required rpm are 50000. Strength of the blade insert and stage's hub should be calculated from this speed and expected blade weights, to verify the capability of materials before fixing engine characteristics. [[Category:Turbofan|Blades]] 4e9acaf1bbc4ee3d022c852933d1c184da95703c 209 208 2011-07-27T23:57:09Z Vincent 1 fixes over the page and turbine section improvements wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. ==Blade fixation== The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. In summary, real engines use a fir tree shape to hold the blade centrifugally, while keeping them free axially. They are then generally axially fixed using a locking screw. This design allows for easy replacement of damaged blades, but is quite complex to manufacture, and requires lots of parts. Since our engine will be smaller, using fixing screws is likely to be not feasible, and a single-spool design will make blades subject to less centrifugal stress. Simpler blade fixation mechanism thus had to be studied. ==Blade designs for efficient air flow== Blades have to be carefully designed, because the overall efficiency of the engine largely depends on their design. '''Stages.''' They are not only made of blades on the rotor, but also blades on the stator, generally called ''vanes''. A '''stage''' is then a pair of a rotor stage and a stator stage, n this order for compressor stages, and in the reverse order for turbine stages. Vanes prevent a rotating air flow to form inside the engine (swirl), driven by the action of the rotor blades. Stator vanes redirect the airflow in the more appropriate direction for the next rotor stage. They increase the energy of the gas ([http://en.wikipedia.org/wiki/Enthalpy enthalpy]) by removing the swirling effect that impairs it. '''Blade shape.''' Most basic design of a fan has flat-shaped blades. Twisted blades with a flat section are an improvement, taking into account the difference in apparent airflow velocity and torque all along the blade. Next step is to have a non-flat section, but an airfoil section. This has been [http://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines proved] to be the only design meeting efficiency requirements of the turbine engines, in 1926 by [http://en.wikipedia.org/wiki/Axial_compressor#Development Alan A. Griffith]. Finally, modern engines are designed with curved edges for the fan, for optimal known efficiency as well as for noise reduction. '''Rotor/stator gaps.''' Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher costs. Anyway, the precision of blades will have to be very good if we don't want the engine to dislocate when it reaches the high rotations-per-minute achieved. A high reproducibility is required and partially automated manufacturing allows it. Design of stages is linked to the energy the blades have to give (compressor and fan) or take (turbine) to the air flow. To better define and understand that energy, we will use standard [http://en.wikipedia.org/wiki/State_function thermodynamic parameters] of gas, a.k.a state variables of a gas, on which are based quantities like the enthalpy: temperature, volume and pressure. We will also use the velocity because the actual work given by a turbofan engine is related to the mass flow rate of the gas expelled by the engine, which relates to velocity of this gas and the state variables. ===Gas variables: temperature, pressure, velocity=== ===Fan design=== The main goal of the fan is to increase the mass flow rate, by increasing the velocity. The mass flow is related to the area of the fan blades and the angular speed of the fan. To increase the velocity, the simplest way is to reduce volume, in other words create a nozzle. The fan duct will thus have to act as a compressor on the aft-end. On the front-end, it is generally designed as an expander, to increase the pressure, allowing more efficient work on the air flow. ===Compressor design=== The main goal of the compressor is to increase pressure. A compressor stage is composed of a blade-mounted rotor and a vane-mounted stator. The shape of the blade is an airfoil and as the gas flows through the rotor and the stator it gains static pressure, since the blades form a expander. Speed is however gained in the rotor section, because of the high rotational speed of the blades. Stator vanes remove the resulting swirl, and converts the velocity (dynamic pressure) to static pressure, thus increasing again the pressure. The volume occupied by the gas can consequently drop as the pressure increases, intake area will be greater than compressor discharge area. Due to friction and pressure rise, temperature will be increased too. [[File:Compressor_blades.png|360px|center]] ===Turbine design=== The main goal of the turbine is to extract energy from the hot and fast gas discharged by the combustion into mechanical (rotational) work. Turbine vanes generally reduce the discharge area, reducing temperature and increasing speed of the flow then impacting the tubine blades. Pressure and temperature may remain constant ([http://en.wikipedia.org/wiki/Turbine#Theory_of_operation impulse turbine]) or not ([http://en.wikipedia.org/wiki/Turbine#Theory_of_operation reaction turbine]) through the turbine rotor. Recent engines feature a reaction turbine. Sectional area decreasing over the turbine blade help energy to be extracted as rotational work while the flow speed increases. ==Mechanical constraints== Blades on all three parts of a turbofan engine undergo heavy mechanical constraints due to high rpm achieved by the rotor, the high temperature in the turbine section and non-negligible temperature in the end of the compressor section, and the high pressure of the gas on which work is performed. Fan blades are not made of plain metal in real engines. In the eighties, they were made in honeycomb composite sandwich material, they are now made in triangular sandwich structure. For an engine of the size we are targeting, fan blades may be built with plain carbon fiber, like the [http://www.schuebeler-jets.com/index.php?option=com_content&task=view&id=102&Itemid=171 Schuebeler HST] high quality R/C engine. Compressor blades are made of titanium alloys, providing high strength and rigidity at these temperatures. Turbine blades are made of nickel alloys, better sustaining the high temperature, and still at higher strength than steel. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing], as a [http://en.wikipedia.org/wiki/Hot_working hot working] process, is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing hot isostatic pressing] possibly too, as explained on the ''How are made turbine blades'' [http://www.youtube.com/watch?v=vN3_Wkyl5PQ video]. I believe that a hot forging press can be done cheaply considering the small size of our blades. For the main fan, it thus may not be used. [[File:Blade.jpg|600px|center]] ''The above picture lacks the airfoil section of the blade, because the CAD software we use doesn't currently support it.'' Work-hardening and heat-resistance processing have to be studied, and depend on the crafting method used in the first place. The high-pressure turbine blades have to face very high temperature and pressure. On real engines, they are made of titanium and nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. Since the required lifetime is lower in our case, we may achieve a working engine with cheaper metals, like steel or nickel-rich alloys for the turbine blades. Historically, the temperature that turbine blades could sustain greatly improved over time, as well as engine efficiency. We can take these two factors and design a reasonably short-lived (6hrs, 3 missions?) and medium efficiency engine. Material for compressor blade and fan blade can probably be aluminum or aluminum alloys like 2024 (dural) or even 7075. The blade insert will have to be milled at some point, unless if it is casted or pressed. Sonic speed at the tip of a 160mm fan is achieved at around 40000 rpm, for 120mm a fan, required rpm are 50000. Strength of the blade insert and stage's hub should be calculated from this speed and expected blade weights, to verify the capability of materials before fixing engine characteristics. [[Category:Turbofan|Blades]] 8f924c4348fe7ba77cc75dbaeaf2fec4869e5943 0 11 206 196 2011-05-09T23:34:49Z Vincent 1 /* Compressor and turbine blades */ fixing ref to Griffin and text on compressor airfoils wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets. They are however very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced while still having a reasonable efficiency, which is our primary concern here. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage feeding the combustion. The fan provides thrust from creating a massive air flow, and the turbine creates thrust by evacuating a hotter but less important air flow. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost has the fan mounted on. One stage engines exist and are less complicated and expensive to build, but are also less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas, and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down because of the price. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing could be to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. Moreover, we may think about a turbine-stage mechanism embedded in the stator to try to cool it, which would make it hollow. The main issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine, because the gas generator (combustion) and the compressor can provide more power to the turbine. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, which would decrease the enthalpy of the gas (its internal energy), driven by the action of rotor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between rotor blades and the stator, as well as between the stator blades and the rotor. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. Blade geometric design by itself can reveal complicated. The first turbine engine(s) had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1926, [http://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines Alan A. Griffin] proved that if blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to deserve being built. Airfoils for blade designs allow compressor stages to better increase the static pressure, since they provide an increasing area for the air flow to pass through (= an expander). ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C, but will deform before melting. Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the stator of the turbine, which is, on actual engines, made of a ceramic coated high temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge circulates, enters the blades, and evacuates through drilled holes in the blades (convective cooling and film cooling again). For the rotor blades, the same principle is used, but with air coming from inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air, ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. The rotor speed information would be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, maybe a simple square-section root is enough. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blades-roots are inserted and fixed. Creating a perpendicular root is already a challenge. Rotor's root would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting is '''to be considered'''. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Compressed air (a.k.a. bleed air) from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while maintaining the combustion. [[Category:Turbofan]] fb3005e8bfcf91c2a0db1944b58b74ecc59f6561 207 206 2011-05-09T23:35:22Z Vincent 1 /* Compressor and turbine blades */ typo wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets. They are however very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced while still having a reasonable efficiency, which is our primary concern here. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage feeding the combustion. The fan provides thrust from creating a massive air flow, and the turbine creates thrust by evacuating a hotter but less important air flow. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost has the fan mounted on. One stage engines exist and are less complicated and expensive to build, but are also less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas, and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down because of the price. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing could be to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. Moreover, we may think about a turbine-stage mechanism embedded in the stator to try to cool it, which would make it hollow. The main issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine, because the gas generator (combustion) and the compressor can provide more power to the turbine. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, which would decrease the enthalpy of the gas (its internal energy), driven by the action of rotor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between rotor blades and the stator, as well as between the stator blades and the rotor. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. Blade geometric design by itself can reveal complicated. The first turbine engine(s) had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1926, [http://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines Alan A. Griffith] proved that if blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to deserve being built. Airfoils for blade designs allow compressor stages to better increase the static pressure, since they provide an increasing area for the air flow to pass through (= an expander). ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C, but will deform before melting. Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the stator of the turbine, which is, on actual engines, made of a ceramic coated high temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge circulates, enters the blades, and evacuates through drilled holes in the blades (convective cooling and film cooling again). For the rotor blades, the same principle is used, but with air coming from inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air, ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. The rotor speed information would be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, maybe a simple square-section root is enough. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blades-roots are inserted and fixed. Creating a perpendicular root is already a challenge. Rotor's root would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting is '''to be considered'''. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Compressed air (a.k.a. bleed air) from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while maintaining the combustion. [[Category:Turbofan]] 1b985289ce19276173be50b12621b9ba0cefaf69 210 207 2011-07-28T00:09:52Z Vincent 1 /* Our design propositions */ adding bearing page wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our Design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets. They are however very difficult to manufacture as well as very expensive. On this page, we will explore how costs can be reduced while still having a reasonable efficiency, which is our primary concern here. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that feeds a turbine, which drives the fan and the compression stage feeding the combustion. The fan provides thrust from creating a massive air flow, and the turbine creates thrust by evacuating a hotter but less important air flow. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost has the fan mounted on. One stage engines exist and are less complicated and expensive to build, but are also less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas, and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.fr/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down because of the price. ===Shaped core or shaped shaft?=== An important optimization to reduce cost and complexity of manufacturing could be to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a massively-turned shaft won't change much on its final mass. Moreover, we may think about a turbine-stage mechanism embedded in the stator to try to cool it, which would make it hollow. The main issue is now how to properly fix the blades to it and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compression blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine, because the gas generator (combustion) and the compressor can provide more power to the turbine. The compressor is not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow to form inside the engine, which would decrease the enthalpy of the gas (its internal energy), driven by the action of rotor blades. Stator blades redirect the airflow on the next compression stage in the more appropriate and efficient direction. Highest efficiency is reached in turbofans when gaps are reduced between rotor blades and the stator, as well as between the stator blades and the rotor. As always, good efficiency means good high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. Blade geometric design by itself can reveal complicated. The first turbine engine(s) had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1926, [http://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines Alan A. Griffith] proved that if blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to deserve being built. Airfoils for blade designs allow compressor stages to better increase the static pressure, since they provide an increasing area for the air flow to pass through (= an expander). ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C, but will deform before melting. Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the stator of the turbine, which is, on actual engines, made of a ceramic coated high temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge circulates, enters the blades, and evacuates through drilled holes in the blades (convective cooling and film cooling again). For the rotor blades, the same principle is used, but with air coming from inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air, ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude, but may be too heavy or expensive to put on the real engine. The rotor speed information would be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and remove axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we need. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, maybe a simple square-section root is enough. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blades-roots are inserted and fixed. Creating a perpendicular root is already a challenge. Rotor's root would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting is '''to be considered'''. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Compressed air (a.k.a. bleed air) from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose innovative [[Turbofan:Alternative Designs|alternative turbofan designs]]. Several pages have been created in the Turbofan [[:Category:Turbofan|category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while maintaining the combustion. * [[Turbofan:Bearings|Bearings and cooling]]: high speed rotations require adapted bearings and cooling, which may be reused for rotor and even turbine cooling. [[Category:Turbofan]] 7a75ced3a7f90b72e3bd798c598bf63837d76475 229 210 2011-10-21T00:20:32Z Vincent 1 Text corrections wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets or other engines. They are however very difficult to design and manufacture and are thus very expensive. On this page, we will explore how costs can be reduced while still having adequate capabilities for high altitude flight. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that puts energy into a gas, this energy is extracted by a turbine, and the turbine drives both the fan that provides thrust and the compression stage that feeds the combustion with oxygen. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake, and so on. The fan provides thrust by creating a massive air flow, and the engine's core also creates thrust by evacuating the high-speed hot combustion gas. In commercial turbofan engines, the fan is generally responsible for 90% of the overall thrust. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost, has the fan mounted on it. One stage engines exist and are less complicated and expensive to build, but are much less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan, but this is not an easy task due to this very speed. Commercial engines featuring a gearbox for the turbofan's fan are expected to reach market in 2012. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan, but are not yet optimal. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seeing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down with the cost. ===Shaped core or shaped shaft?=== An obvious but important optimization to reduce cost and complexity of manufacturing is to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to design and produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates, for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a shaft turned in raw metal won't change much on its final mass. Moreover, we may use a turbine-level mechanism embedded in the stator to try to cool it, which would make it hollow. The main mechanical issues are probably how to properly fix the blades on rotor and stator, how to fix the rotor on the stator with little gap, and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compressor blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys] or are ceramic-coated. It's the inability of blades to withstand heat and work that limit the power of the engine. Indeed, around 70% of the gas provided by the compressor is used only for chamber and turbine cooling, instead of using it to burn more fuel and create more thrust. The compressor and the turbine are not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow driven by the action of rotor blades to form inside the engine, which would decrease the energy of the gas. Stator blades or vanes redirect the airflow on the next stage in the more efficient direction. Highest efficiencies are reached in turbofans when gaps are reduced between rotor blades' tip and the stator, as well as between the stator blades' tip and the rotor. As always, good efficiency means high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. The shape of the blade and the parameters of their cascade also affects the efficiency. A small 5 stage supersonic compressor providing the same pressure rise than a 15 stage subsonic compressor is less efficient, but it may be compensated by the higher thrust-to-weight ratio. Blade geometric design is also very complicated. First turbine engines had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1926, [http://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines Alan A. Griffith] proved that if blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to deserve being built. Airfoils for blade designs allow compressor stages to better increase the static pressure since they create an expander, an increasing area for the air flow to pass through. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C so it will be way off-limits, and even below that, it will deform before melting. And should we mention that blade deformation at high-centrifugal loads, caused by the high rpm, is a very good way to blow the engine off? Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the vanes the turbine, which are, on actual engines, made of a ceramic-coated high-temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge or bleed circulates, enters the blades, and evacuates through small holes in the blades (convective cooling and film cooling). For the rotor blades, the same principle is used, but with compressor air passing inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air or a high speed electric engine for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air and compressor stall, re-ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude. The rotor speed information and altimeter may be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and removed axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we aim. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, a simple square-section root is probably enough. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blade-roots are inserted and fixed. Creating a perpendicular blade root is already a challenge. Rotor blades would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting should be considered. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Air bled from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose [[Turbofan:Alternative Designs|alternative turbofan designs]] based on a mixed-flow turbine. Several pages have been created in the [[:Category:Turbofan|Turbofan category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three to five stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while sustaining the combustion. * [[Turbofan:Bearings|Bearings and cooling]]: high speed rotations require adapted bearings and cooling, which may be reused for rotor and even turbine cooling. [[Category:Turbofan]] 0f2723c96ac183bf1fd714b0e430d5a5d4b911c7 232 229 2011-10-23T23:06:56Z Vincent 1 /* Our design propositions */ design procedure wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets or other engines. They are however very difficult to design and manufacture and are thus very expensive. On this page, we will explore how costs can be reduced while still having adequate capabilities for high altitude flight. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that puts energy into a gas, this energy is extracted by a turbine, and the turbine drives both the fan that provides thrust and the compression stage that feeds the combustion with oxygen. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake, and so on. The fan provides thrust by creating a massive air flow, and the engine's core also creates thrust by evacuating the high-speed hot combustion gas. In commercial turbofan engines, the fan is generally responsible for 90% of the overall thrust. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost, has the fan mounted on it. One stage engines exist and are less complicated and expensive to build, but are much less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan, but this is not an easy task due to this very speed. Commercial engines featuring a gearbox for the turbofan's fan are expected to reach market in 2012. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan, but are not yet optimal. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seeing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down with the cost. ===Shaped core or shaped shaft?=== An obvious but important optimization to reduce cost and complexity of manufacturing is to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to design and produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates, for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a shaft turned in raw metal won't change much on its final mass. Moreover, we may use a turbine-level mechanism embedded in the stator to try to cool it, which would make it hollow. The main mechanical issues are probably how to properly fix the blades on rotor and stator, how to fix the rotor on the stator with little gap, and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compressor blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys] or are ceramic-coated. It's the inability of blades to withstand heat and work that limit the power of the engine. Indeed, around 70% of the gas provided by the compressor is used only for chamber and turbine cooling, instead of using it to burn more fuel and create more thrust. The compressor and the turbine are not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow driven by the action of rotor blades to form inside the engine, which would decrease the energy of the gas. Stator blades or vanes redirect the airflow on the next stage in the more efficient direction. Highest efficiencies are reached in turbofans when gaps are reduced between rotor blades' tip and the stator, as well as between the stator blades' tip and the rotor. As always, good efficiency means high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. The shape of the blade and the parameters of their cascade also affects the efficiency. A small 5 stage supersonic compressor providing the same pressure rise than a 15 stage subsonic compressor is less efficient, but it may be compensated by the higher thrust-to-weight ratio. Blade geometric design is also very complicated. First turbine engines had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1926, [http://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines Alan A. Griffith] proved that if blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to deserve being built. Airfoils for blade designs allow compressor stages to better increase the static pressure since they create an expander, an increasing area for the air flow to pass through. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C so it will be way off-limits, and even below that, it will deform before melting. And should we mention that blade deformation at high-centrifugal loads, caused by the high rpm, is a very good way to blow the engine off? Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the vanes the turbine, which are, on actual engines, made of a ceramic-coated high-temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge or bleed circulates, enters the blades, and evacuates through small holes in the blades (convective cooling and film cooling). For the rotor blades, the same principle is used, but with compressor air passing inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air or a high speed electric engine for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air and compressor stall, re-ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude. The rotor speed information and altimeter may be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and removed axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we aim. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, a simple square-section root is probably enough. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blade-roots are inserted and fixed. Creating a perpendicular blade root is already a challenge. Rotor blades would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting should be considered. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Air bled from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose [[Turbofan:Alternative Designs|alternative turbofan designs]] based on a mixed-flow turbine. Several pages have been created in the [[:Category:Turbofan|Turbofan category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three to five stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while sustaining the combustion. * [[Turbofan:Bearings|Bearings and cooling]]: high speed rotations require adapted bearings and cooling, which may be reused for rotor and even turbine cooling. ===Turbofan design procedure=== # evaluate required thrust (from aircraft mass and lift) # calculate required mass flow rate for the fan # fix bypass ratio and fan diameter and rpm, thus giving core diameter # calculate required power to drive the fan alone # evaluate a gross compressor driving power (refined later)<br />&nbsp; # calculate total power that has to be drawn from the turbine (fan + compressor + losses) # calculate mass flow rate for the combustion alone # calculate mass flow rate for cooling chamber and turbine ## calculate mass flow rate for cooling chamber ## evaluate mass flow rate for cooling turbine to add to the latter ## calculate entropy and fluid parameters at combustor discharge (speed, temperature) ## calculate temperature of turbine vanes and blade and check if it is acceptable ## iterate on item 8.1 until temperature is unacceptable # calculate the number of turbine blades and stages required for this power # calculate compressor discharge pressure and pressure ratio # calculate how many compressor stages are required depending on sonic or supersonic blade design and fix design # calculate compressor driving power # iterate on item 6 until total power varies # design blades for all calculated parameters and re-run at item 6, total power may have changed [[Category:Turbofan]] 9d1dc56d1d967836bed54e5c212ce4b41adb195f 0 47 211 2011-07-28T00:10:32Z Vincent 1 page creation wikitext text/x-wiki =Bearings and cooling= [[Category:Turbofan]] 0d96fd71abac883951f79ad2828fe01436f71ea6 212 211 2011-07-28T00:12:06Z Vincent 1 wikitext text/x-wiki =Bearings and cooling= [[Category:Turbofan|Bearing]] [[Category:Turbofan|Cooling]] 14a10aefe53cd117a1abefacdfa5522546f76172 213 212 2011-07-28T00:12:54Z Vincent 1 /* Bearings and cooling */ wikitext text/x-wiki =Bearings and cooling= [[Category:Turbofan|Bearings and cooling]] 22b2e828633b780d3237eed962164c99c1c7dbf0 214 213 2011-07-28T00:29:35Z Vincent 1 page beginning wikitext text/x-wiki =Bearings and cooling= Rotational speed achieved by the engine will be around 40000rpm. At this speed, regular ball bearings may overheat or suffer from a too fast wear. In real engines, bearings are constantly lubricated by an oil bath, which poses problems with regards to oil leaks in other parts of the engine. Seals are thus places close to bearings to prevent leaks, generally carbon leaks. Accessories like oil pumps, pipes, fixations, filters and so on are also required. ==Bearings== ==Cooling with lubricating oil== ==Oil displacement without pumping== [http://en.wikipedia.org/wiki/Screw_conveyor Screw pumping] will be used as a way to move the oil through the engine, i.e. parts that have to be lubricated and cooled. the work of the turbine will thus directly drive the oil pumping without requiring external accessories. However, cooling the oil may require external hardware, and sealing is absolutely required. [[Category:Turbofan|Bearings and cooling]] e14c3b551a1184af181b27a4daac194f9306e513 215 214 2011-07-28T01:11:27Z Vincent 1 /* Oil displacement without pumping */ wikitext text/x-wiki =Bearings and cooling= Rotational speed achieved by the engine will be around 40000rpm. At this speed, regular ball bearings may overheat or suffer from a too fast wear. In real engines, bearings are constantly lubricated by an oil bath, which poses problems with regards to oil leaks in other parts of the engine. Seals are thus places close to bearings to prevent leaks, generally carbon leaks. Accessories like oil pumps, pipes, fixations, filters and so on are also required. ==Bearings== ==Cooling with lubricating oil== ==Oil displacement without external pumping== [http://en.wikipedia.org/wiki/Screw_conveyor Screw pumping] will be used as a way to move the oil through the engine, i.e. parts that have to be lubricated and cooled. the work of the turbine will thus directly drive the oil pumping without requiring external accessories. However, cooling the oil may require external hardware, and sealing is absolutely required. [[Category:Turbofan|Bearings and cooling]] 6e658c0e4a42f448614f5eece674ec2857018598 226 215 2011-10-19T22:08:21Z Vincent 1 lots of things. Introducing magnetic, fluid, and silicon nitride bearings. wikitext text/x-wiki =Bearings and cooling= Rotational speed achieved by the engine will probably be around 40000rpm. At this speed, regular ball bearings may overheat or suffer from a too fast wear. In real engines, bearings are constantly lubricated by an oil bath, which poses problems with regards to oil pressurization and leaks in other parts of the engine. Seals are consequently placed close to bearings to prevent leaks, generally carbon leaks. Accessories like oil pumps, pipes, fixations, filters and so on, are also required. ==Bearings== ===Ball bearings=== Ball or roller bearings are the obvious way to guide rotating parts. They provide good mechanical constraints on the axis orthogonal to the rotation, they are inexpensive and their integration is reasonably simple. [http://en.wikipedia.org/wiki/Silicon_nitride#Bearings Silicon nitride bearings] have lots of improvements over regular metal ball bearings. Balls are more than 60% less heavy, thus having a lower inertia at high speeds, implying a more softer contact with the tracks, allowing longer lifetime or higher reachable speeds. They also require less lubrication. Fortunately, silicon nitride bearings have reached marked with a large production, and are not over-expensive. ===Alternate bearings=== Fluid or magnetic bearings should be considered. They allow much higher rotation speeds and lower friction, but have two main drawbacks. At standby state, they release the hard constrain on moving parts orthogonally to the rotating axis. In reduced-size turbomachinery, where rotor and stator have to be adjusted to tens of microns, it's quite complicated to use those bearings. The second drawback is that they require more external hardware, to pressurize the fluid or to provide magnetic energy. However, magnetic bearings have been demonstrated in [http://books.google.com/books?id=AXtqMugS3TQC&lpg=PP1&pg=PA263#v=onepage&q&f=false this paper]: S. Jana, V. Arun Kumar and M. Ananda. '''5-axes levitation of a rotor towards indigenization of the magnetic bearing technology'''. In ''Air breathing engines and aerospace propulsion: proceedings of NCABE 2004'', november 2004. ==Use of lubricating oil for cooling== ==Oil displacement without external pumping== [http://en.wikipedia.org/wiki/Screw_conveyor Screw pumping] will be used as a way to move the oil through the engine, i.e. parts that have to be lubricated and cooled. the work of the turbine will thus directly drive the oil pumping without requiring external accessories. However, cooling the oil may require external hardware, and sealing is absolutely required. ==External hardware required for lubrication== A cooling device will be required if the oil gets too hot, which is likely. A basic oil-to-air heat exchanger should be sufficient. Sensors will be required too, at least for oil temperature and displacement. Oil temperature may inform about the status of the engine, and with sufficient experiments and modeling can be used to infer turbine temperature. Oil displacement sensor is required to ensure that there is no problem with the oil/cooling flow in the engine and that the measured temperature is not bogus. [[Category:Turbofan|Bearings and cooling]] ac7005e86b1826ae79dd1598a5de0c66ab52427b 227 226 2011-10-19T22:39:52Z Vincent 1 /* Alternate bearings */ referencing instead of listing the ref. wikitext text/x-wiki =Bearings and cooling= Rotational speed achieved by the engine will probably be around 40000rpm. At this speed, regular ball bearings may overheat or suffer from a too fast wear. In real engines, bearings are constantly lubricated by an oil bath, which poses problems with regards to oil pressurization and leaks in other parts of the engine. Seals are consequently placed close to bearings to prevent leaks, generally carbon leaks. Accessories like oil pumps, pipes, fixations, filters and so on, are also required. ==Bearings== ===Ball bearings=== Ball or roller bearings are the obvious way to guide rotating parts. They provide good mechanical constraints on the axis orthogonal to the rotation, they are inexpensive and their integration is reasonably simple. [http://en.wikipedia.org/wiki/Silicon_nitride#Bearings Silicon nitride bearings] have lots of improvements over regular metal ball bearings. Balls are more than 60% less heavy, thus having a lower inertia at high speeds, implying a more softer contact with the tracks, allowing longer lifetime or higher reachable speeds. They also require less lubrication. Fortunately, silicon nitride bearings have reached marked with a large production, and are not over-expensive. ===Alternate bearings=== Fluid or magnetic bearings should be considered. They allow much higher rotation speeds and lower friction, but have two main drawbacks. At standby state, they release the hard constrain on moving parts orthogonally to the rotating axis. In reduced-size turbomachinery, where rotor and stator have to be adjusted to tens of microns, it's quite complicated to use those bearings. The second drawback is that they require more external hardware, to pressurize the fluid or to provide magnetic energy. However, magnetic bearings have been demonstrated in [http://books.google.com/books?id=AXtqMugS3TQC&lpg=PP1&pg=PA263#v=onepage&q&f=false this paper] <ref>S. Jana, V. Arun Kumar and M. Ananda. '''5-axes levitation of a rotor towards indigenization of the magnetic bearing technology'''. In ''Air breathing engines and aerospace propulsion: proceedings of NCABE 2004'', november 2004.</ref>. ==Use of lubricating oil for cooling== ==Oil displacement without external pumping== [http://en.wikipedia.org/wiki/Screw_conveyor Screw pumping] will be used as a way to move the oil through the engine, i.e. parts that have to be lubricated and cooled. the work of the turbine will thus directly drive the oil pumping without requiring external accessories. However, cooling the oil may require external hardware, and sealing is absolutely required. ==External hardware required for lubrication== A cooling device will be required if the oil gets too hot, which is likely. A basic oil-to-air heat exchanger should be sufficient. Sensors will be required too, at least for oil temperature and displacement. Oil temperature may inform about the status of the engine, and with sufficient experiments and modeling can be used to infer turbine temperature. Oil displacement sensor is required to ensure that there is no problem with the oil/cooling flow in the engine and that the measured temperature is not bogus. [[Category:Turbofan|Bearings and cooling]] 1b548c651b0d3ffa02a798cdd999fd985309d7ea 228 227 2011-10-19T22:40:26Z Vincent 1 Reference section wikitext text/x-wiki =Bearings and cooling= Rotational speed achieved by the engine will probably be around 40000rpm. At this speed, regular ball bearings may overheat or suffer from a too fast wear. In real engines, bearings are constantly lubricated by an oil bath, which poses problems with regards to oil pressurization and leaks in other parts of the engine. Seals are consequently placed close to bearings to prevent leaks, generally carbon leaks. Accessories like oil pumps, pipes, fixations, filters and so on, are also required. ==Bearings== ===Ball bearings=== Ball or roller bearings are the obvious way to guide rotating parts. They provide good mechanical constraints on the axis orthogonal to the rotation, they are inexpensive and their integration is reasonably simple. [http://en.wikipedia.org/wiki/Silicon_nitride#Bearings Silicon nitride bearings] have lots of improvements over regular metal ball bearings. Balls are more than 60% less heavy, thus having a lower inertia at high speeds, implying a more softer contact with the tracks, allowing longer lifetime or higher reachable speeds. They also require less lubrication. Fortunately, silicon nitride bearings have reached marked with a large production, and are not over-expensive. ===Alternate bearings=== Fluid or magnetic bearings should be considered. They allow much higher rotation speeds and lower friction, but have two main drawbacks. At standby state, they release the hard constrain on moving parts orthogonally to the rotating axis. In reduced-size turbomachinery, where rotor and stator have to be adjusted to tens of microns, it's quite complicated to use those bearings. The second drawback is that they require more external hardware, to pressurize the fluid or to provide magnetic energy. However, magnetic bearings have been demonstrated in [http://books.google.com/books?id=AXtqMugS3TQC&lpg=PP1&pg=PA263#v=onepage&q&f=false this paper] <ref>S. Jana, V. Arun Kumar and M. Ananda. '''5-axes levitation of a rotor towards indigenization of the magnetic bearing technology'''. In ''Air breathing engines and aerospace propulsion: proceedings of NCABE 2004'', november 2004.</ref>. ==Use of lubricating oil for cooling== ==Oil displacement without external pumping== [http://en.wikipedia.org/wiki/Screw_conveyor Screw pumping] will be used as a way to move the oil through the engine, i.e. parts that have to be lubricated and cooled. the work of the turbine will thus directly drive the oil pumping without requiring external accessories. However, cooling the oil may require external hardware, and sealing is absolutely required. ==External hardware required for lubrication== A cooling device will be required if the oil gets too hot, which is likely. A basic oil-to-air heat exchanger should be sufficient. Sensors will be required too, at least for oil temperature and displacement. Oil temperature may inform about the status of the engine, and with sufficient experiments and modeling can be used to infer turbine temperature. Oil displacement sensor is required to ensure that there is no problem with the oil/cooling flow in the engine and that the measured temperature is not bogus. ==References== <references /> [[Category:Turbofan|Bearings and cooling]] b13e8360103cad05f18dbaae4fa91d538faa3ce3 230 228 2011-10-23T00:37:42Z Vincent 1 /* Alternate bearings */ miti's work wikitext text/x-wiki =Bearings and cooling= Rotational speed achieved by the engine will probably be around 40000rpm. At this speed, regular ball bearings may overheat or suffer from a too fast wear. In real engines, bearings are constantly lubricated by an oil bath, which poses problems with regards to oil pressurization and leaks in other parts of the engine. Seals are consequently placed close to bearings to prevent leaks, generally carbon leaks. Accessories like oil pumps, pipes, fixations, filters and so on, are also required. ==Bearings== ===Ball bearings=== Ball or roller bearings are the obvious way to guide rotating parts. They provide good mechanical constraints on the axis orthogonal to the rotation, they are inexpensive and their integration is reasonably simple. [http://en.wikipedia.org/wiki/Silicon_nitride#Bearings Silicon nitride bearings] have lots of improvements over regular metal ball bearings. Balls are more than 60% less heavy, thus having a lower inertia at high speeds, implying a more softer contact with the tracks, allowing longer lifetime or higher reachable speeds. They also require less lubrication. Fortunately, silicon nitride bearings have reached marked with a large production, and are not over-expensive. ===Alternate bearings=== Fluid or magnetic bearings should be considered. They allow much higher rotation speeds and lower friction, but have two main drawbacks. At standby state, they release the hard constrain on moving parts orthogonally to the rotating axis. In reduced-size turbomachinery, where rotor and stator have to be adjusted to tens of microns, it's quite complicated to use those bearings. The second drawback is that they require more external hardware, to pressurize the fluid or to provide magnetic energy. However, magnetic bearings have been demonstrated in [http://books.google.com/books?id=AXtqMugS3TQC&lpg=PP1&pg=PA263#v=onepage&q&f=false this paper] <ref>S. Jana, V. Arun Kumar and M. Ananda. '''5-axes levitation of a rotor towards indigenization of the magnetic bearing technology'''. In ''Air breathing engines and aerospace propulsion: proceedings of NCABE 2004'', november 2004.</ref>, in which axial position accuracy is measured below 150µm for a 4kg rotor at around 2000rpm. The rotor position sensor has a resolution of 2µm per mV. However, no indication is given about the resting position of the rotor and how that impacts the clearance between rotor and stator. [http://en.wikipedia.org/wiki/Foil_bearing Foil bearings] are a particular type of fluid bearing, that "Unlike aero or hydrostatic bearings, foil bearings require no external pressurisation system for the working fluid, so the hydrodynamic bearing is self-starting". In [http://b-dig.iie.org.mx/BibDig/P06-0351/pdfs/track-16/GT2006-90791.pdf this other paper] <ref>Hooshang Heshmat, Michael J. Tomaszewski, James F. Walton II. '''Small gas turbine engine operating with high temperature foil bearing'''. In ''proceedeings of GT2006 ASME Turbo Expo 2006: Power for land, sea and air'', may 2006.</ref>, a small centrifugal turbojet is built to evaluate the ability of [http://www.miti.cc/products-services.html MiTi]'s product, a foil bearing, to sustain very high rotation speeds (120'000rpm) and high temperature (800°C). The bearing has a low spacing between the rotor's journal and the stator fixation, but it is secured, in this paper, using a ball bearing on the compressor side, where the temperature is low. They planned to make a dual-foil bearing, we'll need to check on that. MiTi also demonstrated a [http://www.miti.cc/newsletters/20_150mm_foil_journal_bearing%20_hybrid_foil_magnetic_bearing.pdf hybrid foil magnetic bearing], that has the advantages of magnetic bearings at low speeds and those of foil bearings at high speeds. ==Use of lubricating oil for cooling== ==Oil displacement without external pumping== [http://en.wikipedia.org/wiki/Screw_conveyor Screw pumping] will be used as a way to move the oil through the engine, i.e. parts that have to be lubricated and cooled. the work of the turbine will thus directly drive the oil pumping without requiring external accessories. However, cooling the oil may require external hardware, and sealing is absolutely required. ==External hardware required for lubrication== A cooling device will be required if the oil gets too hot, which is likely. A basic oil-to-air heat exchanger should be sufficient. Sensors will be required too, at least for oil temperature and displacement. Oil temperature may inform about the status of the engine, and with sufficient experiments and modeling can be used to infer turbine temperature. Oil displacement sensor is required to ensure that there is no problem with the oil/cooling flow in the engine and that the measured temperature is not bogus. ==References== <references /> [[Category:Turbofan|Bearings and cooling]] 5c10a33cc20e966aa6530b6b0fc7279fc8059d6f 0 1 216 125 2011-10-09T19:11:49Z Vincent 1 /* N-Prize and low-cost space access reflections */ title change wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This web site aims to gather to gather my researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official web site for the [[N-Prize]]. The official web site is here: http://www.n-prize.com/ . The goal of the competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. I'm not part of a team, nor did I register a team, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2011. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. a4257d555353537b2a0113afc21943d1283d8ab4 217 216 2011-10-09T23:17:50Z Vincent 1 /* N-Prize and reflections on low-cost access to space */ link to founder and guidelines, otaski first appearance too wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative (OTASKI), currently being founded. I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2012. Anyway, if you find this project interesting, you can still [[Join and guidelines|join and participate]]! ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 3a79049931b7a07e4e684e93991f89ae120a7c98 219 217 2011-10-09T23:23:28Z Vincent 1 /* N-Prize and reflections on low-cost access to space */ different links for join and participate wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative (OTASKI), currently being founded. I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2012. Anyway, if you find this project interesting, you can still [[Join|join]] and [[Guidelines|participate]]! ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. c047b434892ebb2303e97ad192b057b1e17d4d0c 221 219 2011-10-17T00:10:35Z Vincent 1 adding the current status section wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative (OTASKI), currently being founded. I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2012. Anyway, if you find this project interesting, you can still [[Join|join]] and [[Guidelines|participate]]! ==Current status of the project== Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. Since the main constraint is to have low costs, we'll have to build the plane first, including the turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus focus on the turbofan research and design. The first step is to have the engine designed and build. We may consider helping other N-Prize teams if this is done in time, or other projects outside the contest, by providing them those engines and help with aircraft design. Some other parts of the aircraft/rocket are currently being studied, for example the software control and the low-cost sensors that can be used to provide autonomy to the aircraft at first, then make the rocket go into space and reach orbit. ''(updated october 2011)'' Information available on this site is quite outdated, and may reflect some false information, since it was done with little knowledge about fluid dynamics and turbines. Time is a hard resource to manage, and we will probably build a documentation base somehow to provide access to all information used to develop the project, and update the website pages so that they really reflect current state. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. d62d15dbe3eb209ed9fb6b4bee9f7b67dc8f732b 222 221 2011-10-18T00:45:18Z Vincent 1 /* Current status of the project */ links wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative (OTASKI), currently being founded. I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2012. Anyway, if you find this project interesting, you can still [[Join|join]] and [[Guidelines|participate]]! ==Current status of the project== Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. Since the main constraint is to have low costs, we'll have to build the plane first, including the turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus focus on the turbofan research and design. The first step is to have the engine [[Build_a_cheap_turbofan|designed]] and built. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are currently being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. ''(updated october 2011)'' Information available on this site is quite outdated, and may reflect some weak or false information, since it was done with little knowledge on the topics. Time is a hard resource to manage, and we will probably build a documentation base somehow to provide access to all information used to develop the project, and update the website pages so that they really reflect the progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 9f00c2e883857387ce73a1eaec65162b0ee3aede 247 222 2011-11-15T02:28:37Z Vincent 1 /* Current status of the project */ wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative (OTASKI), currently being founded. I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2012. Anyway, if you find this project interesting, you can still [[Join|join]] and [[Guidelines|participate]]! ==Current status of the project / News== Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. Since the main constraint is to have low costs, we'll have to build the plane first, including the turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus focus on the turbofan research and design. The first step is to have the engine [[Build_a_cheap_turbofan|designed]] and built. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are currently being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. '''''(updated November 2011)''''' Information available on this site is sometimes quite outdated, and may reflect some weak or false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 4816185aa098a7cbc42f987673609b293f33a998 248 247 2011-11-16T12:06:35Z Vincent 1 /* N-Prize and reflections on low-cost access to space */ otaski update wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2012. Anyway, if you find this project interesting, you can still [[Join|join]] and [[Guidelines|participate]]! Maybe with several people we can still make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==Current status of the project / News== Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. Since the main constraint is to have low costs, we'll have to build the plane first, including the turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus focus on the turbofan research and design. The first step is to have the engine [[Build_a_cheap_turbofan|designed]] and built. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are currently being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. '''''(updated November 2011)''''' Information available on this site is sometimes quite outdated, and may reflect some weak or false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is low speed compared to orbital speed though. Can electricity energy be considered for that kind of mission ? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, the main issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] ===Staging and recovery=== Separation from the rocket is a big concern. If wings are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. ===Guidance=== GPS can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. adf9b2b12b782874ee946ced0e0bd7e0eef5adc8 0 48 218 2011-10-09T23:21:11Z Vincent 1 page creation, no info. wikitext text/x-wiki =Founder of the project= No information available on the founder yet. 48e1a730e6ffa0eeaec62c4400ba53fe342def3a 250 218 2011-11-17T03:32:22Z Vincent 1 Redirected page to [[User:Vincent]] wikitext text/x-wiki #REDIRECT [[User:Vincent]] 7109c3fb44145d60311837de03799ba6483643ae 0 49 220 2011-10-09T23:37:32Z Vincent 1 page creation wikitext text/x-wiki =Join the project!= This project is willing to provide easy access to complicated information. Technologies are studied and compared, and results are clearly explained in order to explain why we have to chose one solution instead of another. Innovative solutions are also proposed and evaluated. All ressources are made available freely, in accordance with OTASKI, check the licence here or there. ==Why join?== Lots of technologies and scientific fields of research are covered by aerospatial activity. Help is welcome, because it would take too much years to understand deeply all the details of each field. If you can provide information, ressources or if you can orientate the choices made for the prototypes, [[Founder|I]]'d be glad to have you the project. ==How to join?== Start by creating an account. Then, explain on your user page here or by email to join@nprize.mine.nu what you would like to do in the project. That's all. 55d27b90246d7169dcf2fa48f869d54671156a6b 249 220 2011-11-16T12:14:46Z Vincent 1 /* How to join? */ account request wikitext text/x-wiki =Join the project!= This project is willing to provide easy access to complicated information. Technologies are studied and compared, and results are clearly explained in order to explain why we have to chose one solution instead of another. Innovative solutions are also proposed and evaluated. All ressources are made available freely, in accordance with OTASKI, check the licence here or there. ==Why join?== Lots of technologies and scientific fields of research are covered by aerospatial activity. Help is welcome, because it would take too much years to understand deeply all the details of each field. If you can provide information, ressources or if you can orientate the choices made for the prototypes, [[Founder|I]]'d be glad to have you the project. ==How to join?== [[Special:RequestAccount|Creating an account]] on the wiki allows you to participate to the project. It has to be accepted by administrators before you can use it. Account requests have to be completed with a minimum of information about what you would like to do in the project or in what domain you have some expertise. Thank you! ef4b9de1b33525e3662f18fe7c6124ba42f64377 0 16 223 181 2011-10-18T01:01:24Z Vincent 1 link fixes wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/l amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific parts:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. 930a0e96f96b4f4dd8ab625c76dfd72265c5a11a 224 223 2011-10-19T00:04:12Z Vincent 1 /* Specific parts: */ wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/l amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. 3971acffda021cdd937a14606750eb1a78802aac 225 224 2011-10-19T00:29:59Z Vincent 1 /* Web pages */ adding two lectures and the section wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/l amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. ebe31a0afce5b4ec9a3376e1c8e7a93be5a1df31 237 225 2011-11-01T22:06:23Z Vincent 1 /* Turbines and turbofans */ klaus97 wikitext text/x-wiki =Resources= This page gathers all documentation available on the numerous subjects linked to rocket science (and turbofans). ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/l amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] by Klaus Hünecke, 1997. * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. 9bab08ea1fe8572af3484d8519fbcc561116ca37 0 30 231 197 2011-10-23T01:04:08Z Vincent 1 /* Compressor design */ outdate warning and text fixes wikitext text/x-wiki =Compressor design= '''Pictures on this page are outdated and do not reflect the actual compressor design'''. Real-world engines have 10 to 20 compression stages. We will first design a 3- to 5-stage compressor, in a single-spool engine and assess if more stages are required (= more expensive). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. The two pictures below show an overall sketching of the compressor with all its blades, flat rendered, but will be updated to be airfoil rendered someday if we can find how to draw it. Both rotor and stator will have an expander and reaction design. [[File:Compressor_noshell.jpg|300px]] [[File:Compressor_side.jpg|300px]] The three stages depicted here have the following number of blades (same for rotor and stator of the same stage): * Stage1: 18 blades (20° spaced) * Stage2: 20 blades (18° spaced) * Stage3: 24 blades (15° spaced) ==Rotor design== The following pictures represent the shaft and the three stages compressor rotor blades. [[File:Rotor_front_side.jpg|300px]] [[File:Rotor_side.jpg|300px]] Special manufacturing and balancing care have to be given to the rotor, since it will spin at very high rotation speeds. A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's first stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. This design is depicted below. [[File:Blade_fixation_simple.jpg|300px|center]] An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This keyed design is also considered for higher stages, and is illustrated in the models below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade insert and the hub: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more usual solution to this problem, based on a real standalone hub in which are milled the inserts holes. Blades' inserts would remain a linear part and the milling of the hub would thus be linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly and easily fixed. A keyed design could also be added on the extremities of the blades, i.e. the side of the hub. The pictures below represent this mounting system, in exploded view (larger versions available). The hub is the grey round part with only one insert milled in it, the yellow part is the inter-stage spacer that has the key holding the blades. [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the blades mounted on the hub (in orange) and on the key system from the inter-stage spacer (in yellow). [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== The stator will have as main issue the fact that blades must have a very thin insert. Shaping them like a T should be considered, but bolting them on the stator looks more promising for now. [[File:Stator_side.jpg|300px]] [[File:Stator_front.jpg|300px]] [[Category:Turbofan|Compressor]] 82b942035bbc4638d7ab62cc86dcc25bb655e4ea 233 231 2011-10-23T23:40:40Z Vincent 1 /* Compressor design */ generally accepted compressor properties wikitext text/x-wiki =Compressor design= '''Pictures on this page are outdated and do not reflect the actual compressor design'''. Several properties of axial compressors should be considered: * a constant energy rise should happen between each stage. It is sometimes said that the temperature should have a constant rise, instead of the energy (entropy, based on both pressure and temperature). The direct implication of those two concerns is that each stage has to have a lower compression ratio than the previous, since its input flow properties have already been risen. The first stage will thus have the highest pressure rise. * the axial air flow velocity in the compressor should be constant. Since dynamic pressure is traded for static pressure, if the cross-section area is kept constant, the flow speed falls. That's why the area containing the flow in the compressor decreases along stages. * maximum pressure rise for a subsonic compressor is around 1.6:1. Since we tend towards a singe-spool engine, and the fan should not be several times supersonic, the compressor is likely to be subsonic. From these three concerns, a 4 stages compressor with the following stage pressure ratios: 1.55, 1.52, 1.47, 1.41; would have a pressure ratio of 4.88, which is terrible. A fifth stage with 1.35 PR would make a final PR of 6.59... Real-world engines have 10 to 20 compression stages. We will first design a 3- to 5-stage compressor, in a single-spool engine and assess if more stages are required (= more expensive). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. The two pictures below show an overall sketching of the compressor with all its blades, flat rendered, but will be updated to be airfoil rendered someday if we can find how to draw it. Both rotor and stator will have an expander and reaction design. [[File:Compressor_noshell.jpg|300px]] [[File:Compressor_side.jpg|300px]] The three stages depicted here have the following number of blades (same for rotor and stator of the same stage): * Stage1: 18 blades (20° spaced) * Stage2: 20 blades (18° spaced) * Stage3: 24 blades (15° spaced) ==Rotor design== The following pictures represent the shaft and the three stages compressor rotor blades. [[File:Rotor_front_side.jpg|300px]] [[File:Rotor_side.jpg|300px]] Special manufacturing and balancing care have to be given to the rotor, since it will spin at very high rotation speeds. A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's first stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. This design is depicted below. [[File:Blade_fixation_simple.jpg|300px|center]] An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This keyed design is also considered for higher stages, and is illustrated in the models below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade insert and the hub: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more usual solution to this problem, based on a real standalone hub in which are milled the inserts holes. Blades' inserts would remain a linear part and the milling of the hub would thus be linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly and easily fixed. A keyed design could also be added on the extremities of the blades, i.e. the side of the hub. The pictures below represent this mounting system, in exploded view (larger versions available). The hub is the grey round part with only one insert milled in it, the yellow part is the inter-stage spacer that has the key holding the blades. [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the blades mounted on the hub (in orange) and on the key system from the inter-stage spacer (in yellow). [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== The stator will have as main issue the fact that blades must have a very thin insert. Shaping them like a T should be considered, but bolting them on the stator looks more promising for now. [[File:Stator_side.jpg|300px]] [[File:Stator_front.jpg|300px]] [[Category:Turbofan|Compressor]] 49ef34ae9f2a6b0e6cc8d6709306a041e9efb146 0 43 234 191 2011-10-31T21:28:08Z Vincent 1 vbpr wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine behind it, and their two flows mix at the exhaust, inside the engine for low bypass ratio engines and outside for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a quite hot gas flow, which would eventually corrode or fatigue it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans, at least while research or people won't have proven it was wrong: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. The hybrid turbine is a mix of axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its course, while extracting some of its energy for shaft rotation work. The aft-fan would intake the mixed flow of the fresh intake and the turbine discharge, providing higher energy to the fan flow. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000- series aluminum alloys. Besides, the resulting design will inevitably be a longer engine. ==Full transonic engine design in a single spool with 2.0 ''virtual'' BPR== We speak here of ''virtual'' BPR because since flows are mixed before the fan intake, there is no clear separation between flows of the fan and the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). ''Drawings (schematics or 3D CAD models) are coming soon.'' [[Category:Turbofan|Alternative design]] d9a80ed5e70fffddec87a5401ffc95aa1b872acc 235 234 2011-10-31T23:35:37Z Vincent 1 transonic design wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine behind it, and their two flows mix at the exhaust, inside the engine for low bypass ratio engines and outside for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a quite hot gas flow, which would eventually corrode or fatigue it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans, at least while research or people won't have proven it was wrong: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. The hybrid turbine is a mix of axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its course, while extracting some of its energy for shaft rotation work. The aft-fan would intake the mixed flow of the fresh intake and the turbine discharge, providing higher energy to the fan flow. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000- series aluminum alloys. Besides, the resulting design will inevitably be a longer engine. ==Full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' BPR because since flows are mixed before the fan intake, there is no clear separation between flows of the fan and the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is actually the sum of both. Let's take a example turbofan engine with a 13cm fan and a 9cm core (and compressor) diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both side, and the BPR is in fact 2.09. A great advantage of our aft-fan engine design is that both the first stage(s) of the compressor and the fan can operate at transonic speeds. A blade or a fan is said having a transonic operation when the flow passing through it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, while their root is generally subsonic too. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound which depends on the temperature of the air flow. A transonic operation allows higher compression ratios to be achieved, with a lower efficiency than subsonic operation as a drawback. As a result, our design allows both the compressor entry stage and the fan to operate at transonic speeds, while they have the same rotation speed (one spool shaft), at high altitudes. This is possible because of the temperature difference in air passing through both. The compressor will breathe fresh air, which can go down to -57°C in altitude. In this case, the sonic speed will be 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. For a 200°C rise, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. Having a transonic speed operation allows higher the compression ratios. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor, it's not able to have a high overall compression ratio. With a transonic-operating first stage, we hope to have at least 2.0 CR for it, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Efficiency of the fan will also allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. ''Drawings (schematics or 3D CAD models) are coming soon.'' [[Category:Turbofan|Alternative design]] 2ca5bab65649ce37e5609aa5e2d8336181996893 236 235 2011-11-01T00:19:26Z Vincent 1 /* Full transonic engine design in a single spool with 2.1 BPR */ reading again and corrections wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine behind it, and their two flows mix at the exhaust, inside the engine for low bypass ratio engines and outside for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a quite hot gas flow, which would eventually corrode or fatigue it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans, at least while research or people won't have proven it was wrong: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. The hybrid turbine is a mix of axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its course, while extracting some of its energy for shaft rotation work. The aft-fan would intake the mixed flow of the fresh intake and the turbine discharge, providing higher energy to the fan flow. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000- series aluminum alloys. Besides, the resulting design will inevitably be a longer engine. ==Full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' bypass ratio (BPR) because since flows are mixed before the fan intake, there is no clear separation between flows of the fan and of the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is thus the ratio between the sum of the fan duct mass flow rate and the core's mass flow rate over the core's mass flow rate (M_fan + M_core) / M_core. Let's take an example turbofan engine with a 13cm fan and a 9cm core (and compressor) diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both side ((13²-9²)/9²), and the BPR is in fact 2.09 (13²/9²). A great advantage of our aft-fan engine design is that both the first stage(s) of the compressor and the fan can operate at '''transonic''' speeds. A blade or a fan is said having a [http://en.wikipedia.org/wiki/Transonic transonic] operation when the flow passing through it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, while their root is generally subsonic too. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound, which depends on the temperature of the air flow. A transonic operation allows higher compression ratios to be achieved, with a lower efficiency than subsonic operation as a drawback, around 5% less because of the drag induced by shock waves. As a result, our design allows both the compressor entry stage and the fan to operate at transonic speeds, while they have the same rotation speed (one spool shaft), at high altitudes. This is possible because of the temperature difference in air passing through both. The compressor will breathe fresh air, which can go down as low as -57°C in altitude. In this case, the sonic speed is 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. If we fix a 200°C rise of this air flow, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. Having a transonic speed operation allows higher the compression ratios. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor, it's not able to have a high overall compression ratio (CR). With a transonic-operating first stage, we hope to have at least 2.0 CR for it, instead of the maximum 1.6 CR in subsonic operation, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Higher efficiency of the fan will allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. ''Drawings (schematics or 3D CAD models) are coming soon.'' [[Category:Turbofan|Alternative design]] 478052d2760a1b78cd92a02742c741d14e278274 242 236 2011-11-02T20:00:44Z Vincent 1 /* Full transonic engine design in a single spool with 2.1 BPR */ images wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine behind it, and their two flows mix at the exhaust, inside the engine for low bypass ratio engines and outside for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a quite hot gas flow, which would eventually corrode or fatigue it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans, at least while research or people won't have proven it was wrong: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. The hybrid turbine is a mix of axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its course, while extracting some of its energy for shaft rotation work. The aft-fan would intake the mixed flow of the fresh intake and the turbine discharge, providing higher energy to the fan flow. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000- series aluminum alloys. Besides, the resulting design will inevitably be a longer engine. ==Full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' bypass ratio (BPR) because since flows are mixed before the fan intake, there is no clear separation between flows of the fan and of the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is thus the ratio between the sum of the fan duct mass flow rate and the core's mass flow rate over the core's mass flow rate (M_fan + M_core) / M_core. Let's take an example turbofan engine with a 13cm fan and a 9cm core (and compressor) diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both side ((13²-9²)/9²), and the BPR is in fact 2.09 (13²/9²). A great advantage of our aft-fan engine design is that both the first stage(s) of the compressor and the fan can operate at '''transonic''' speeds. A blade or a fan is said having a [http://en.wikipedia.org/wiki/Transonic transonic] operation when the flow passing through it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, while their root is generally subsonic too. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound, which depends on the temperature of the air flow. A transonic operation allows higher compression ratios to be achieved, with a lower efficiency than subsonic operation as a drawback, around 5% less because of the drag induced by shock waves. As a result, our design allows both the compressor entry stage and the fan to operate at transonic speeds, while they have the same rotation speed (one spool shaft), at high altitudes. This is possible because of the temperature difference in air passing through both. The compressor will breathe fresh air, which can go down as low as -57°C in altitude. In this case, the sonic speed is 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. If we fix a 200°C rise of this air flow, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. Having a transonic speed operation allows higher the compression ratios. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor, it's not able to have a high overall compression ratio (CR). With a transonic-operating first stage, we hope to have at least 2.0 CR for it, instead of the maximum 1.6 CR in subsonic operation, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Higher efficiency of the fan will allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. [[File:Engine core and fan.jpg|300px]] [[File:Engine core and fan side.jpg|350px]] ''The above schematics may not reflect actual proportions and have transparency inconsistencies.'' That said, the engine's core is the green internal tube. The fan sketched inside it, at the front of the engine, is the compressor's first stage rotor. The stator and other stage are not represented, and will obviously be placed on the conic part behind it. The red part is the annular [[Turbofan:Combustors|combustor]], not easy to see clearly in these pictures. '''The turbine is not represented''' either, but an axial rotor should be placed right after the combustor, and the hybrid turbine will be milled in the shaft in the curved part after a second axial turbine guide vane. We can see that the discharge of the turbine is mixed with the inlet air of the aft-mounted fan. The fan duct is thus longer than in traditional front-mounted engines, but the nozzle can be closer to the fan, allowing to work on both core and fan flow. [[Category:Turbofan|Alternative design]] 2b428c8bcf2c42228529e1079418a054c99d5eb3 0 44 238 195 2011-11-01T22:29:46Z Vincent 1 first draft wikitext text/x-wiki =Turbofan combustor= ==Typical designs== A good summary can be found at chapter 5 of the [[Resources#Turbines_and_turbofans|book]] "Jet engines: fundamentals of theory, design, and operation". We summarize it even more here. Three main designs exist: can combustors, can-annular combustor, and annular combustor. They were invented in that order, to solve some of the flaws of previous models. The basic principle is that the combustor takes the compressor discharge flow, which has a high velocity, a high pressure, and a high enough temperature for the combustion to occur with a reduced risk of flaming-out. The first job of the combustor is to slow down the flow so that it can be mixed to fuel and provide a stable flame. The second task of the combustor is to prevent the flame from spreading outside it, and thus preventing a too hot flow to spread on other parts of the engine. Its design ensures that the flame will be properly confined, and that the output gas temperature will be acceptable for the turbine. The hotter the gas is discharged in the turbine, the more energy will be gained from the combustion. To limit this temperature, and also to limit the combustor temperature to avoid its own melting down, diluting air is introduced in the combustor. A can combustor features mechanisms to slow down the input air flow, fuel injection nozzle, air dilution flows, and a hot gas discharge. Several can combustors are then placed circumferentially in the engine's core. The main disadvantage is the loss of useful volume in the combustor section, and the non-uniformity of the hot gas discharge. The can-annular combustor has a single air dilution diffuser for all combustion cans, making more efficient use of the volume and reducing the design complexity and weight. Finally, annular combustors are designed as a single annular part taking the maximum available volume from the section. All modern engines use this design, so do we. ==Our design== Flame holding, Cooling, Material, Ignition. [[Category:Turbofan|Combustors]] f237852ec2482548d018123270b2b1ee2447c488 239 238 2011-11-01T22:38:05Z Vincent 1 flow properties wikitext text/x-wiki =Turbofan combustor= ==Typical designs== A good summary can be found at chapter 5 of the [[Resources#Turbines_and_turbofans|book]] "Jet engines: fundamentals of theory, design, and operation". We summarize it even more here. Three main designs exist: can combustors, can-annular combustor, and annular combustor. They were invented in that order, to solve some of the flaws of previous models. The basic principle is that the combustor takes the compressor discharge flow, which has a high velocity, a high pressure, and a high enough temperature for the combustion to occur with a reduced risk of flaming-out. The first job of the combustor is to slow down the flow so that it can be mixed to fuel and provide a stable flame. The second task of the combustor is to prevent the flame from spreading outside it, and thus preventing a too hot flow to spread on other parts of the engine. Its design ensures that the flame will be properly confined, and that the output gas temperature will be acceptable for the turbine. The hotter the gas is discharged in the turbine, the more energy will be gained from the combustion. To limit this temperature, and also to limit the combustor temperature to avoid its own melting down, diluting air is introduced in the combustor. A can combustor features mechanisms to slow down the input air flow, fuel injection nozzle, air dilution flows, and a hot gas discharge. Several can combustors are then placed circumferentially in the engine's core. The main disadvantage is the loss of useful volume in the combustor section, and the non-uniformity of the hot gas discharge. The can-annular combustor has a single air dilution diffuser for all combustion cans, making more efficient use of the volume and reducing the design complexity and weight. Finally, annular combustors are designed as a single annular part taking the maximum available volume from the section. All modern engines use this design, so do we. ==Combustor properties== Combustor, like other parts of the engine, operate with air flow. It thus affects its properties, like temperature, pressure and velocity. A combustor is said efficient when it minimizes the pressure drop. The air used for combustion must be slowed down, and this is generally done by both expanding the input flow and swirling a part of it. This swirl causes turbulences that imply pressure loss, and thus efficiency reduction of the combustor. The produced hot gas will be used to provide energy to the engine, and this energy needs to be in an easily extractable form. High temperatures and high velocities are appreciable for turbine work extraction. There is probably no upper bound for velocity, but there is one for temperature, the temperature that the turbine inlet vanes can sustain, and the temperature that the turbine blades can sustain. ==Our design== Flame holding, Cooling, Material, Ignition. [[Category:Turbofan|Combustors]] 6487bbb97a517f1070bedcbde2a984ccee4968ba 244 239 2011-11-02T20:13:47Z Vincent 1 /* Our design */ sketch of the combustor wikitext text/x-wiki =Turbofan combustor= ==Typical designs== A good summary can be found at chapter 5 of the [[Resources#Turbines_and_turbofans|book]] "Jet engines: fundamentals of theory, design, and operation". We summarize it even more here. Three main designs exist: can combustors, can-annular combustor, and annular combustor. They were invented in that order, to solve some of the flaws of previous models. The basic principle is that the combustor takes the compressor discharge flow, which has a high velocity, a high pressure, and a high enough temperature for the combustion to occur with a reduced risk of flaming-out. The first job of the combustor is to slow down the flow so that it can be mixed to fuel and provide a stable flame. The second task of the combustor is to prevent the flame from spreading outside it, and thus preventing a too hot flow to spread on other parts of the engine. Its design ensures that the flame will be properly confined, and that the output gas temperature will be acceptable for the turbine. The hotter the gas is discharged in the turbine, the more energy will be gained from the combustion. To limit this temperature, and also to limit the combustor temperature to avoid its own melting down, diluting air is introduced in the combustor. A can combustor features mechanisms to slow down the input air flow, fuel injection nozzle, air dilution flows, and a hot gas discharge. Several can combustors are then placed circumferentially in the engine's core. The main disadvantage is the loss of useful volume in the combustor section, and the non-uniformity of the hot gas discharge. The can-annular combustor has a single air dilution diffuser for all combustion cans, making more efficient use of the volume and reducing the design complexity and weight. Finally, annular combustors are designed as a single annular part taking the maximum available volume from the section. All modern engines use this design, so do we. ==Combustor properties== Combustor, like other parts of the engine, operate with air flow. It thus affects its properties, like temperature, pressure and velocity. A combustor is said efficient when it minimizes the pressure drop. The air used for combustion must be slowed down, and this is generally done by both expanding the input flow and swirling a part of it. This swirl causes turbulences that imply pressure loss, and thus efficiency reduction of the combustor. The produced hot gas will be used to provide energy to the engine, and this energy needs to be in an easily extractable form. High temperatures and high velocities are appreciable for turbine work extraction. There is probably no upper bound for velocity, but there is one for temperature, the temperature that the turbine inlet vanes can sustain, and the temperature that the turbine blades can sustain. ==Our design== The following sketch represents the position and the shape of the combustor on the shaft (below in green) and the core's shell (above and in the front of the picture with transparency). All air will be forced into the combustor, mainly at the middle and the rear part of it to dilute the hot gas. Its section increases at front, to allow burning to happen, and decreases at rear, which will speed up and cool down exit gases too. [[File:Combustor_sketch_section.jpg|350px]] ''TODO: Flame holding, Cooling, Material, Ignition.'' [[Category:Turbofan|Combustors]] 00a10f0684fbcbf9ed6b996bf8ec389842c903a6 245 244 2011-11-03T23:41:49Z Vincent 1 /* Typical designs */ text fixes wikitext text/x-wiki =Turbofan combustor= ==Typical designs== A good summary can be found at chapter 5 of the [[Resources#Turbines_and_turbofans|book]] "Jet engines: fundamentals of theory, design, and operation". We summarize it even more here. Three main designs exist: can combustors, can-annular combustor, and annular combustor. They were invented in that order, to solve some of the flaws of previous models. The basic principle is that the combustor takes the compressor discharge flow, which has a high velocity, a high pressure, and a high enough temperature for the combustion to occur with a reduced risk of flaming-out. The first job of the combustor is to slow down the flow so that it can be mixed to fuel and provide a stable flame. The second task of the combustor is to prevent the flame from spreading outside it, and thus preventing a too hot flow to spread on other parts of the engine. Its design ensures that the flame will be properly confined, and that the output gas temperature will be acceptable for the turbine. The hotter the gas is discharged in the turbine, the more energy will be gained from the combustion. To limit this temperature, and also to limit the combustor temperature to avoid its own melting down, diluting air is introduced in the combustor. A can combustor features all mechanisms required for a proper combustion: input air flow slow down, fuel injection nozzles, air dilution flows, and a hot gas discharge. Several cans are then placed circumferentially in the engine's core. The main drawback is the loss of useful volume in the combustor section, and the non-uniformity of the hot gas discharge. The can-annular combustor has removed the external shell of cans, allowing a single air dilution diffuser to exist for all combustion cans, which is actually the core's shell itself. They make more efficient use of the volume and reducing the design complexity and weight. Finally, annular combustors are designed as a single annular part taking all the available volume from the combustion section. All modern engines use this design, so do we. ==Combustor properties== Combustor, like other parts of the engine, operate with air flow. It thus affects its properties, like temperature, pressure and velocity. A combustor is said efficient when it minimizes the pressure drop. The air used for combustion must be slowed down, and this is generally done by both expanding the input flow and swirling a part of it. This swirl causes turbulences that imply pressure loss, and thus efficiency reduction of the combustor. The produced hot gas will be used to provide energy to the engine, and this energy needs to be in an easily extractable form. High temperatures and high velocities are appreciable for turbine work extraction. There is probably no upper bound for velocity, but there is one for temperature, the temperature that the turbine inlet vanes can sustain, and the temperature that the turbine blades can sustain. ==Our design== The following sketch represents the position and the shape of the combustor on the shaft (below in green) and the core's shell (above and in the front of the picture with transparency). All air will be forced into the combustor, mainly at the middle and the rear part of it to dilute the hot gas. Its section increases at front, to allow burning to happen, and decreases at rear, which will speed up and cool down exit gases too. [[File:Combustor_sketch_section.jpg|350px]] ''TODO: Flame holding, Cooling, Material, Ignition.'' [[Category:Turbofan|Combustors]] 4bdfed0ba8ad637a47b155023f1f210aa3be78aa 246 245 2011-11-04T00:58:38Z Vincent 1 /* Our design */ sections draft wikitext text/x-wiki =Turbofan combustor= ==Typical designs== A good summary can be found at chapter 5 of the [[Resources#Turbines_and_turbofans|book]] "Jet engines: fundamentals of theory, design, and operation". We summarize it even more here. Three main designs exist: can combustors, can-annular combustor, and annular combustor. They were invented in that order, to solve some of the flaws of previous models. The basic principle is that the combustor takes the compressor discharge flow, which has a high velocity, a high pressure, and a high enough temperature for the combustion to occur with a reduced risk of flaming-out. The first job of the combustor is to slow down the flow so that it can be mixed to fuel and provide a stable flame. The second task of the combustor is to prevent the flame from spreading outside it, and thus preventing a too hot flow to spread on other parts of the engine. Its design ensures that the flame will be properly confined, and that the output gas temperature will be acceptable for the turbine. The hotter the gas is discharged in the turbine, the more energy will be gained from the combustion. To limit this temperature, and also to limit the combustor temperature to avoid its own melting down, diluting air is introduced in the combustor. A can combustor features all mechanisms required for a proper combustion: input air flow slow down, fuel injection nozzles, air dilution flows, and a hot gas discharge. Several cans are then placed circumferentially in the engine's core. The main drawback is the loss of useful volume in the combustor section, and the non-uniformity of the hot gas discharge. The can-annular combustor has removed the external shell of cans, allowing a single air dilution diffuser to exist for all combustion cans, which is actually the core's shell itself. They make more efficient use of the volume and reducing the design complexity and weight. Finally, annular combustors are designed as a single annular part taking all the available volume from the combustion section. All modern engines use this design, so do we. ==Combustor properties== Combustor, like other parts of the engine, operate with air flow. It thus affects its properties, like temperature, pressure and velocity. A combustor is said efficient when it minimizes the pressure drop. The air used for combustion must be slowed down, and this is generally done by both expanding the input flow and swirling a part of it. This swirl causes turbulences that imply pressure loss, and thus efficiency reduction of the combustor. The produced hot gas will be used to provide energy to the engine, and this energy needs to be in an easily extractable form. High temperatures and high velocities are appreciable for turbine work extraction. There is probably no upper bound for velocity, but there is one for temperature, the temperature that the turbine inlet vanes can sustain, and the temperature that the turbine blades can sustain. ==Our design== The following sketch represents the position and the shape of the combustor on the shaft (below in green) and the core's shell (above and in the front of the picture with transparency). All air will be forced into the combustor, mainly at the middle and the rear part of it to dilute the hot gas. Its section increases at front, to allow burning to happen, and decreases at rear, which will speed up and cool down exit gases too. [[File:Combustor_sketch_section.jpg|350px]] ===Flame keeping=== In a classical engine, the flame can be sustained not too hardly because there is always enough air to feed it. In a reduced size engine, the available air is obviously lower, so is the fuel mass flow rate. With lower amount of fuel or air flow disruptions, the mixture may become hardly flammable. Careful nozzle design and air flow paths will have to be studied. Two approaches are common to flame keeping. In combustor first, the air is slowed down by expanding it and making it pass through a hole labyrinth and finally a swirler, the latter being also required for proper mixing of the fuel and air. Second, in an afterburner, where the flow velocity is much higher, a turbulence-making device is placed on the hot gas path and the fuel is injected and igniter in the turbulence. The turbulence device is often simply a pipe with a cross-section in form of a '<', the angle being pointed against the flow. ===Injector nozzles=== The nozzle will dictate the diffusion volume of the fuel in the combustor. Mixing efficiency with air depends on it, and temperature of combustor walls too. Indeed, if the volume, generally a cone, is too wide, the flame will touch the walls or prevent dilution air to enter as expected. If it's too narrow, the flame will be harder to sustain on flow disruptions. ===Ignition=== Ignition will be provided by a RC model glow plug, if ignition tests succeed (unlikely). Otherwise, a spark igniter will be created and a single spark device will be placed in the combustor. If we were confident about the engine, we could also ignite it when the plane is on the ground, with an external device like compressed air device for engine startup. But with first tests, flame-outs are likely to occur. It would be more safe to have the sparker on board. Besides, the usefulness of the on board igniter should be assessed when there's no startup device (compressed air, DC motor...) on board. [[Category:Turbofan|Combustors]] 041903f4d2d91555ea4672ca3e1ca8fbf25575ef 6 50 240 2011-11-02T19:35:01Z Vincent 1 Sketch of the engine, with the core and the fan duct, aft-mounted fan and first stage compressor rotor drafted. The chamber is in red/brown. wikitext text/x-wiki Sketch of the engine, with the core and the fan duct, aft-mounted fan and first stage compressor rotor drafted. The chamber is in red/brown. 684a5601e60a21d9fbc5063068ebf1206b285d3e 6 51 241 2011-11-02T19:36:51Z Vincent 1 Sketch of the engine, with the core and the fan duct, aft-mounted fan and first stage compressor rotor drafted. The chamber is in red/brown. wikitext text/x-wiki Sketch of the engine, with the core and the fan duct, aft-mounted fan and first stage compressor rotor drafted. The chamber is in red/brown. 684a5601e60a21d9fbc5063068ebf1206b285d3e 6 52 243 2011-11-02T20:08:49Z Vincent 1 Sketch of the combustor cross-section. The green part is the shaft, the transparent light blue part is the core shell. wikitext text/x-wiki Sketch of the combustor cross-section. The green part is the shaft, the transparent light blue part is the core shell. 78b7ecf8997b758c7d69542e24a6f0286ceeddf1 2 6 251 15 2011-11-17T03:35:03Z Vincent 1 little user page update wikitext text/x-wiki =Founder of the project= Greetings, voyager from the stars! I, Vincent, am the founder of this project and the administrator of this web site. 815914bff22e562baaf0875b2cdb39a8f3017973 270 251 2012-02-09T01:05:12Z Vincent 1 /* Founder of the project */ wikitext text/x-wiki =Founder of the project= Greetings, voyager from the stars! I, Vincent, am the founder of this project and the administrator of this web site. I won't put my email address here to avoid spam, but if you try vincent at the host name of this website it should work, or you can also reach me by requesting an account (link at the top-right of the page). 8399b764dd82f50a26407794c902d8f6ab0ef041 293 270 2012-05-27T15:20:05Z Vincent 1 wikitext text/x-wiki =Founder of the project= Greetings! I, Vincent, am the founder of this project and the administrator of this web site. I won't put my email address here to avoid spam, but if you try vincent at the host name of this website it should work, or you can also reach me by requesting an account, the link is at the top-right of all pages. 25ffb76c9e5439dc5b433b5f5adde4d6a8750c46 0 30 252 233 2011-11-27T01:28:58Z Vincent 1 adding broyce link wikitext text/x-wiki =Compressor design= '''Pictures on this page are outdated and do not reflect the actual compressor design'''. Several properties of axial compressors should be considered: * a constant energy rise should happen between each stage. It is sometimes said that the temperature should have a constant rise, instead of the energy (entropy, based on both pressure and temperature). The direct implication of those two concerns is that each stage has to have a lower compression ratio than the previous, since its input flow properties have already been risen. The first stage will thus have the highest pressure rise. * the axial air flow velocity in the compressor should be constant. Since dynamic pressure is traded for static pressure, if the cross-section area is kept constant, the flow speed falls. That's why the area containing the flow in the compressor decreases along stages. * maximum pressure rise for a subsonic compressor is around 1.6:1. Since we tend towards a singe-spool engine, and the fan should not be several times supersonic, the compressor is likely to be subsonic. From these three concerns, a 4 stages compressor with the following stage pressure ratios: 1.55, 1.52, 1.47, 1.41; would have a pressure ratio of 4.88, which is terrible. A fifth stage with 1.35 PR would make a final PR of 6.59... Real-world engines have 10 to 20 compression stages. We will first design a 3- to 5-stage compressor, in a single-spool engine and assess if more stages are required (= more expensive). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. The two pictures below show an overall sketching of the compressor with all its blades, flat rendered, but will be updated to be airfoil rendered someday if we can find how to draw it. Both rotor and stator will have an expander and reaction design. [[File:Compressor_noshell.jpg|300px]] [[File:Compressor_side.jpg|300px]] The three stages depicted here have the following number of blades (same for rotor and stator of the same stage): * Stage1: 18 blades (20° spaced) * Stage2: 20 blades (18° spaced) * Stage3: 24 blades (15° spaced) ==Rotor design== The following pictures represent the shaft and the three stages compressor rotor blades. [[File:Rotor_front_side.jpg|300px]] [[File:Rotor_side.jpg|300px]] Special manufacturing and balancing care have to be given to the rotor, since it will spin at very high rotation speeds. A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's first stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. This design is depicted below. [[File:Blade_fixation_simple.jpg|300px|center]] An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This keyed design is also considered for higher stages, and is illustrated in the models below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade insert and the hub: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more usual solution to this problem, based on a real standalone hub in which are milled the inserts holes. Blades' inserts would remain a linear part and the milling of the hub would thus be linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly and easily fixed. A keyed design could also be added on the extremities of the blades, i.e. the side of the hub. The pictures below represent this mounting system, in exploded view (larger versions available). The hub is the grey round part with only one insert milled in it, the yellow part is the inter-stage spacer that has the key holding the blades. [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the blades mounted on the hub (in orange) and on the key system from the inter-stage spacer (in yellow). [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== The stator will have as main issue the fact that blades must have a very thin insert. Shaping them like a T should be considered, but bolting them on the stator looks more promising for now. [[File:Stator_side.jpg|300px]] [[File:Stator_front.jpg|300px]] ==Resources and references== * The chapter "Axial-Flow Compressors" from Meherwan P. Broyce is a nice way to get started with axial compressor design ([http://www.netl.doe.gov/technologies/coalpower/turbines/refshelf/handbook/2.0.pdf pdf]). [[Category:Turbofan|Compressor]] 25f3fd6c4d7212da20482f9c24228c33ea8ee8e4 257 252 2011-11-27T02:23:50Z Vincent 1 adding velocity triangles wikitext text/x-wiki =Compressor design= '''Pictures on this page are outdated and do not reflect the actual compressor design (text is not up-to-date either)'''. Several properties of axial compressors should be considered: * a constant energy rise should happen between each stage. It is sometimes said that the temperature should have a constant rise, instead of the energy (entropy, based on both pressure and temperature). The direct implication of those two concerns is that each stage has to have a lower compression ratio than the previous, since its input flow properties have already been risen. The first stage will thus have the highest pressure rise. * the axial air flow velocity in the compressor should be constant. Since dynamic pressure is traded for static pressure, if the cross-section area is kept constant, the flow speed falls. That's why the area containing the flow in the compressor decreases along stages. * maximum pressure rise for a '''subsonic''' compressor is around 1.6:1. Since we tend towards a singe-spool engine, and the fan should not be several times supersonic, the compressor is likely to be subsonic. From these three concerns, a 4 stages compressor with the following stage pressure ratios: 1.55, 1.52, 1.47, 1.41; would have a pressure ratio of 4.88, which is terrible. A fifth stage with 1.35 PR would make a final PR of 6.59... Real-world engines have 10 to 20 compression stages. We will first design a 3- to 5-stage compressor, in a single-spool engine and assess if more stages are required (= more expensive). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. The two pictures below show an overall sketching of the compressor with all its blades, flat rendered, but will be updated to be airfoil rendered someday if we can find how to draw it. Both rotor and stator will have an expander and reaction design. [[File:Compressor_noshell.jpg|300px]] [[File:Compressor_side.jpg|300px]] The three stages depicted here have the following number of blades (same for rotor and stator of the same stage): * Stage1: 18 blades (20° spaced) * Stage2: 20 blades (18° spaced) * Stage3: 24 blades (15° spaced) ==Blade designs== Blade design is closely related to the rotating speed, especially their incidence and deviation angle. The following picture is well known in the compressor or turbine blading world and depicts the velocity triangle. As blades rotate, they both undergo the air flow of the compressor and of their rotation's apparent wind. The triangle allows to calculate the effective air flow direction at rotor input and discharge, and to align stator blades accordingly. [[File:Velocity_triangles.png|300px|center]] ==Rotor design== The following pictures represent the shaft and the three stages compressor rotor blades. [[File:Rotor_front_side.jpg|300px]] [[File:Rotor_side.jpg|300px]] Special manufacturing and balancing care have to be given to the rotor, since it will spin at very high rotation speeds. A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's first stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. This design is depicted below. [[File:Blade_fixation_simple.jpg|300px|center]] An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This keyed design is also considered for higher stages, and is illustrated in the models below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade insert and the hub: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more usual solution to this problem, based on a real standalone hub in which are milled the inserts holes. Blades' inserts would remain a linear part and the milling of the hub would thus be linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly and easily fixed. A keyed design could also be added on the extremities of the blades, i.e. the side of the hub. The pictures below represent this mounting system, in exploded view (larger versions available). The hub is the grey round part with only one insert milled in it, the yellow part is the inter-stage spacer that has the key holding the blades. [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the blades mounted on the hub (in orange) and on the key system from the inter-stage spacer (in yellow). [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== The stator will have as main issue the fact that blades must have a very thin insert. Shaping them like a T should be considered, but bolting them on the stator looks more promising for now. [[File:Stator_side.jpg|300px]] [[File:Stator_front.jpg|300px]] ==Resources and references== * The chapter "Axial-Flow Compressors" from Meherwan P. Broyce is a nice way to get started with axial compressor design ([http://www.netl.doe.gov/technologies/coalpower/turbines/refshelf/handbook/2.0.pdf pdf]). [[Category:Turbofan|Compressor]] 64af6cc7bae3533cbca8d09977ca558bf47a8176 258 257 2011-11-27T03:47:56Z Vincent 1 /* Resources and references */ aerofoil analysis programs wikitext text/x-wiki =Compressor design= '''Pictures on this page are outdated and do not reflect the actual compressor design (text is not up-to-date either)'''. Several properties of axial compressors should be considered: * a constant energy rise should happen between each stage. It is sometimes said that the temperature should have a constant rise, instead of the energy (entropy, based on both pressure and temperature). The direct implication of those two concerns is that each stage has to have a lower compression ratio than the previous, since its input flow properties have already been risen. The first stage will thus have the highest pressure rise. * the axial air flow velocity in the compressor should be constant. Since dynamic pressure is traded for static pressure, if the cross-section area is kept constant, the flow speed falls. That's why the area containing the flow in the compressor decreases along stages. * maximum pressure rise for a '''subsonic''' compressor is around 1.6:1. Since we tend towards a singe-spool engine, and the fan should not be several times supersonic, the compressor is likely to be subsonic. From these three concerns, a 4 stages compressor with the following stage pressure ratios: 1.55, 1.52, 1.47, 1.41; would have a pressure ratio of 4.88, which is terrible. A fifth stage with 1.35 PR would make a final PR of 6.59... Real-world engines have 10 to 20 compression stages. We will first design a 3- to 5-stage compressor, in a single-spool engine and assess if more stages are required (= more expensive). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. The two pictures below show an overall sketching of the compressor with all its blades, flat rendered, but will be updated to be airfoil rendered someday if we can find how to draw it. Both rotor and stator will have an expander and reaction design. [[File:Compressor_noshell.jpg|300px]] [[File:Compressor_side.jpg|300px]] The three stages depicted here have the following number of blades (same for rotor and stator of the same stage): * Stage1: 18 blades (20° spaced) * Stage2: 20 blades (18° spaced) * Stage3: 24 blades (15° spaced) ==Blade designs== Blade design is closely related to the rotating speed, especially their incidence and deviation angle. The following picture is well known in the compressor or turbine blading world and depicts the velocity triangle. As blades rotate, they both undergo the air flow of the compressor and of their rotation's apparent wind. The triangle allows to calculate the effective air flow direction at rotor input and discharge, and to align stator blades accordingly. [[File:Velocity_triangles.png|300px|center]] ==Rotor design== The following pictures represent the shaft and the three stages compressor rotor blades. [[File:Rotor_front_side.jpg|300px]] [[File:Rotor_side.jpg|300px]] Special manufacturing and balancing care have to be given to the rotor, since it will spin at very high rotation speeds. A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's first stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. This design is depicted below. [[File:Blade_fixation_simple.jpg|300px|center]] An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This keyed design is also considered for higher stages, and is illustrated in the models below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade insert and the hub: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more usual solution to this problem, based on a real standalone hub in which are milled the inserts holes. Blades' inserts would remain a linear part and the milling of the hub would thus be linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly and easily fixed. A keyed design could also be added on the extremities of the blades, i.e. the side of the hub. The pictures below represent this mounting system, in exploded view (larger versions available). The hub is the grey round part with only one insert milled in it, the yellow part is the inter-stage spacer that has the key holding the blades. [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the blades mounted on the hub (in orange) and on the key system from the inter-stage spacer (in yellow). [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== The stator will have as main issue the fact that blades must have a very thin insert. Shaping them like a T should be considered, but bolting them on the stator looks more promising for now. [[File:Stator_side.jpg|300px]] [[File:Stator_front.jpg|300px]] ==Resources and references== ===Documentation on axial flow compressors=== * The chapter "Axial-Flow Compressors" from Meherwan P. Broyce is a nice way to get started with axial compressor design ([http://www.netl.doe.gov/technologies/coalpower/turbines/refshelf/handbook/2.0.pdf pdf]). ===Aerofoil analysis programs=== * A list of existing programs for aerofoil analysis and design is available [http://forums.x-plane.org/index.php?showtopic=54036 here]. * Free and quality software from that list seem to be [http://web.mit.edu/drela/Public/web/xfoil/ XFOIL] and [http://www.desktop.aero/panda.php PANDA]. [[Category:Turbofan|Compressor]] 6f2c8e51b62e4d819afd171cf181ae20180db0ee 260 258 2011-11-28T00:47:08Z Vincent 1 /* Resources and references */ adding notes on propulsion wikitext text/x-wiki =Compressor design= '''Pictures on this page are outdated and do not reflect the actual compressor design (text is not up-to-date either)'''. Several properties of axial compressors should be considered: * a constant energy rise should happen between each stage. It is sometimes said that the temperature should have a constant rise, instead of the energy (entropy, based on both pressure and temperature). The direct implication of those two concerns is that each stage has to have a lower compression ratio than the previous, since its input flow properties have already been risen. The first stage will thus have the highest pressure rise. * the axial air flow velocity in the compressor should be constant. Since dynamic pressure is traded for static pressure, if the cross-section area is kept constant, the flow speed falls. That's why the area containing the flow in the compressor decreases along stages. * maximum pressure rise for a '''subsonic''' compressor is around 1.6:1. Since we tend towards a singe-spool engine, and the fan should not be several times supersonic, the compressor is likely to be subsonic. From these three concerns, a 4 stages compressor with the following stage pressure ratios: 1.55, 1.52, 1.47, 1.41; would have a pressure ratio of 4.88, which is terrible. A fifth stage with 1.35 PR would make a final PR of 6.59... Real-world engines have 10 to 20 compression stages. We will first design a 3- to 5-stage compressor, in a single-spool engine and assess if more stages are required (= more expensive). [[Turbofan:Blades|Blade design and manufacturing]] have a dedicated page. The two pictures below show an overall sketching of the compressor with all its blades, flat rendered, but will be updated to be airfoil rendered someday if we can find how to draw it. Both rotor and stator will have an expander and reaction design. [[File:Compressor_noshell.jpg|300px]] [[File:Compressor_side.jpg|300px]] The three stages depicted here have the following number of blades (same for rotor and stator of the same stage): * Stage1: 18 blades (20° spaced) * Stage2: 20 blades (18° spaced) * Stage3: 24 blades (15° spaced) ==Blade designs== Blade design is closely related to the rotating speed, especially their incidence and deviation angle. The following picture is well known in the compressor or turbine blading world and depicts the velocity triangle. As blades rotate, they both undergo the air flow of the compressor and of their rotation's apparent wind. The triangle allows to calculate the effective air flow direction at rotor input and discharge, and to align stator blades accordingly. [[File:Velocity_triangles.png|300px|center]] ==Rotor design== The following pictures represent the shaft and the three stages compressor rotor blades. [[File:Rotor_front_side.jpg|300px]] [[File:Rotor_side.jpg|300px]] Special manufacturing and balancing care have to be given to the rotor, since it will spin at very high rotation speeds. A blade mounting failure will likely cause the loss of the aircraft, given the constraints. The part that holds the blades is called the '''hub'''. We'll call the part of the blades that is hold by the blades at their root the '''insert'''. ===Fan and compressor's first stage blades mounting=== For the first compressor stage and the fan, blades are long enough to have a root parallel to the shaft. An innovative design for the hub would be to a use spacer part between each blade. The whole compressor stage fixation would be the result of all blades and spacer put together side by side. It would be like a pie chart, in which separations are the blades' inserts. This design is probably not used on real planes for two reasons: it's more heavy, since the inside of the compressor stage/shaft is full of metal, and it's too complicated for higher stages as we will see below. This design is depicted below. [[File:Blade_fixation_simple.jpg|300px|center]] An alternative design would be to have the blades moving freely from the hub, and when assembled, a key of some sort would fix them on the hub. The hub and the blades would have a notch in the same place, in which a ring-shaped key would place to hold everything together. This keyed design is also considered for higher stages, and is illustrated in the models below. ===2nd and 3rd stage's compressor blade close-up and manufacturing=== Higher stages have shorter blades, and need to have the base of the blade non-parallel to the shaft for their twisting to be acceptable. That brings problems to the design and manufacturing of the blade insert and the hub: creating a part that can be duplicated to form a cylinder, but with a non-parallel-to-the-axis separation. We propose a more usual solution to this problem, based on a real standalone hub in which are milled the inserts holes. Blades' inserts would remain a linear part and the milling of the hub would thus be linear too. Since it's a round part, this makes a strange effect, but allows the blades to be properly and easily fixed. A keyed design could also be added on the extremities of the blades, i.e. the side of the hub. The pictures below represent this mounting system, in exploded view (larger versions available). The hub is the grey round part with only one insert milled in it, the yellow part is the inter-stage spacer that has the key holding the blades. [[File:Interblade1.jpg|300px]] [[File:Interblade2.jpg|300px]] Below is a close-up on the blades mounted on the hub (in orange) and on the key system from the inter-stage spacer (in yellow). [[File:Blade_fixation1.jpg|300px]] [[File:Blade_fixation2.jpg|300px]] ==Stator design== The stator will have as main issue the fact that blades must have a very thin insert. Shaping them like a T should be considered, but bolting them on the stator looks more promising for now. [[File:Stator_side.jpg|300px]] [[File:Stator_front.jpg|300px]] ==Resources and references== ===Documentation on axial flow compressors=== * The chapter "Axial-Flow Compressors" from Meherwan P. Broyce is a nice way to get started with axial compressor design ([http://www.netl.doe.gov/technologies/coalpower/turbines/refshelf/handbook/2.0.pdf pdf]). * Notes from the School of Aerospace Engineering (Georgia Institute of Technology) AE4451 Propulsion, Winter/ Spring 2002 ([http://www.adl.gatech.edu/classes/propulsion/prop12.html html]). ===Aerofoil analysis programs=== * A list of existing programs for aerofoil analysis and design is available [http://forums.x-plane.org/index.php?showtopic=54036 here]. * Free and quality software from that list seem to be [http://web.mit.edu/drela/Public/web/xfoil/ XFOIL] and [http://www.desktop.aero/panda.php PANDA]. [[Category:Turbofan|Compressor]] f60bf83a6e268608fb4a5eb56a74527832d72a0f 6 53 253 2011-11-27T01:29:33Z Vincent 1 Nomenclature of aerofoils (used in fan/compressor/turbine blades). Source is the chapter "2.0 Axial-Flow Compressors" from M. P. Broyce. wikitext text/x-wiki Nomenclature of aerofoils (used in fan/compressor/turbine blades). Source is the chapter "2.0 Axial-Flow Compressors" from M. P. Broyce. c869f708b9bf44473e5e63639789af4662178a15 0 32 254 209 2011-11-27T01:32:47Z Vincent 1 /* Blade designs for efficient air flow */ adding nomenclature wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. ==Blade fixation== The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. In summary, real engines use a fir tree shape to hold the blade centrifugally, while keeping them free axially. They are then generally axially fixed using a locking screw. This design allows for easy replacement of damaged blades, but is quite complex to manufacture, and requires lots of parts. Since our engine will be smaller, using fixing screws is likely to be not feasible, and a single-spool design will make blades subject to less centrifugal stress. Simpler blade fixation mechanism thus had to be studied. ==Blade designs for efficient air flow== Blades have to be carefully designed, because the overall efficiency of the engine largely depends on their design. '''Stages.''' They are not only made of blades on the rotor, but also blades on the stator, generally called ''vanes''. A '''stage''' is then a pair of a rotor stage and a stator stage, n this order for compressor stages, and in the reverse order for turbine stages. Vanes prevent a rotating air flow to form inside the engine (swirl), driven by the action of the rotor blades. Stator vanes redirect the airflow in the more appropriate direction for the next rotor stage. They increase the energy of the gas ([http://en.wikipedia.org/wiki/Enthalpy enthalpy]) by removing the swirling effect that impairs it. '''Blade shape.''' Most basic design of a fan has flat-shaped blades. Twisted blades with a flat section are an improvement, taking into account the difference in apparent airflow velocity and torque all along the blade. Next step is to have a non-flat section, but an airfoil section. This has been [http://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines proved] to be the only design meeting efficiency requirements of the turbine engines, in 1926 by [http://en.wikipedia.org/wiki/Axial_compressor#Development Alan A. Griffith]. Finally, modern engines are designed with curved edges for the fan, for optimal known efficiency as well as for noise reduction. '''Rotor/stator gaps.''' Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher costs. Anyway, the precision of blades will have to be very good if we don't want the engine to dislocate when it reaches the high rotations-per-minute achieved. A high reproducibility is required and partially automated manufacturing allows it. Design of stages is linked to the energy the blades have to give (compressor and fan) or take (turbine) to the air flow. To better define and understand that energy, we will use standard [http://en.wikipedia.org/wiki/State_function thermodynamic parameters] of gas, a.k.a state variables of a gas, on which are based quantities like the enthalpy: temperature, volume and pressure. We will also use the velocity because the actual work given by a turbofan engine is related to the mass flow rate of the gas expelled by the engine, which relates to velocity of this gas and the state variables. ===Gas variables: temperature, pressure, velocity=== ===Nomenclature=== Click on the following image to see it larger. All main terms are depicted. [[File:Blade_nomenclature.png|360px|center]] ===Fan design=== The main goal of the fan is to increase the mass flow rate, by increasing the velocity. The mass flow is related to the area of the fan blades and the angular speed of the fan. To increase the velocity, the simplest way is to reduce volume, in other words create a nozzle. The fan duct will thus have to act as a compressor on the aft-end. On the front-end, it is generally designed as an expander, to increase the pressure, allowing more efficient work on the air flow. ===Compressor design=== The main goal of the compressor is to increase pressure. A compressor stage is composed of a blade-mounted rotor and a vane-mounted stator. The shape of the blade is an airfoil and as the gas flows through the rotor and the stator it gains static pressure, since the blades form a expander. Speed is however gained in the rotor section, because of the high rotational speed of the blades. Stator vanes remove the resulting swirl, and converts the velocity (dynamic pressure) to static pressure, thus increasing again the pressure. The volume occupied by the gas can consequently drop as the pressure increases, intake area will be greater than compressor discharge area. Due to friction and pressure rise, temperature will be increased too. [[File:Compressor_blades.png|360px|center]] ===Turbine design=== The main goal of the turbine is to extract energy from the hot and fast gas discharged by the combustion into mechanical (rotational) work. Turbine vanes generally reduce the discharge area, reducing temperature and increasing speed of the flow then impacting the tubine blades. Pressure and temperature may remain constant ([http://en.wikipedia.org/wiki/Turbine#Theory_of_operation impulse turbine]) or not ([http://en.wikipedia.org/wiki/Turbine#Theory_of_operation reaction turbine]) through the turbine rotor. Recent engines feature a reaction turbine. Sectional area decreasing over the turbine blade help energy to be extracted as rotational work while the flow speed increases. ==Mechanical constraints== Blades on all three parts of a turbofan engine undergo heavy mechanical constraints due to high rpm achieved by the rotor, the high temperature in the turbine section and non-negligible temperature in the end of the compressor section, and the high pressure of the gas on which work is performed. Fan blades are not made of plain metal in real engines. In the eighties, they were made in honeycomb composite sandwich material, they are now made in triangular sandwich structure. For an engine of the size we are targeting, fan blades may be built with plain carbon fiber, like the [http://www.schuebeler-jets.com/index.php?option=com_content&task=view&id=102&Itemid=171 Schuebeler HST] high quality R/C engine. Compressor blades are made of titanium alloys, providing high strength and rigidity at these temperatures. Turbine blades are made of nickel alloys, better sustaining the high temperature, and still at higher strength than steel. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing], as a [http://en.wikipedia.org/wiki/Hot_working hot working] process, is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing hot isostatic pressing] possibly too, as explained on the ''How are made turbine blades'' [http://www.youtube.com/watch?v=vN3_Wkyl5PQ video]. I believe that a hot forging press can be done cheaply considering the small size of our blades. For the main fan, it thus may not be used. [[File:Blade.jpg|600px|center]] ''The above picture lacks the airfoil section of the blade, because the CAD software we use doesn't currently support it.'' Work-hardening and heat-resistance processing have to be studied, and depend on the crafting method used in the first place. The high-pressure turbine blades have to face very high temperature and pressure. On real engines, they are made of titanium and nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. Since the required lifetime is lower in our case, we may achieve a working engine with cheaper metals, like steel or nickel-rich alloys for the turbine blades. Historically, the temperature that turbine blades could sustain greatly improved over time, as well as engine efficiency. We can take these two factors and design a reasonably short-lived (6hrs, 3 missions?) and medium efficiency engine. Material for compressor blade and fan blade can probably be aluminum or aluminum alloys like 2024 (dural) or even 7075. The blade insert will have to be milled at some point, unless if it is casted or pressed. Sonic speed at the tip of a 160mm fan is achieved at around 40000 rpm, for 120mm a fan, required rpm are 50000. Strength of the blade insert and stage's hub should be calculated from this speed and expected blade weights, to verify the capability of materials before fixing engine characteristics. [[Category:Turbofan|Blades]] 119492514ddfe430df9e2702427dc42f289ff3d8 256 254 2011-11-27T02:06:55Z Vincent 1 /* Compressor design */ link to main article for compressor wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. ==Blade fixation== The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. In summary, real engines use a fir tree shape to hold the blade centrifugally, while keeping them free axially. They are then generally axially fixed using a locking screw. This design allows for easy replacement of damaged blades, but is quite complex to manufacture, and requires lots of parts. Since our engine will be smaller, using fixing screws is likely to be not feasible, and a single-spool design will make blades subject to less centrifugal stress. Simpler blade fixation mechanism thus had to be studied. ==Blade designs for efficient air flow== Blades have to be carefully designed, because the overall efficiency of the engine largely depends on their design. '''Stages.''' They are not only made of blades on the rotor, but also blades on the stator, generally called ''vanes''. A '''stage''' is then a pair of a rotor stage and a stator stage, n this order for compressor stages, and in the reverse order for turbine stages. Vanes prevent a rotating air flow to form inside the engine (swirl), driven by the action of the rotor blades. Stator vanes redirect the airflow in the more appropriate direction for the next rotor stage. They increase the energy of the gas ([http://en.wikipedia.org/wiki/Enthalpy enthalpy]) by removing the swirling effect that impairs it. '''Blade shape.''' Most basic design of a fan has flat-shaped blades. Twisted blades with a flat section are an improvement, taking into account the difference in apparent airflow velocity and torque all along the blade. Next step is to have a non-flat section, but an airfoil section. This has been [http://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines proved] to be the only design meeting efficiency requirements of the turbine engines, in 1926 by [http://en.wikipedia.org/wiki/Axial_compressor#Development Alan A. Griffith]. Finally, modern engines are designed with curved edges for the fan, for optimal known efficiency as well as for noise reduction. '''Rotor/stator gaps.''' Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher costs. Anyway, the precision of blades will have to be very good if we don't want the engine to dislocate when it reaches the high rotations-per-minute achieved. A high reproducibility is required and partially automated manufacturing allows it. Design of stages is linked to the energy the blades have to give (compressor and fan) or take (turbine) to the air flow. To better define and understand that energy, we will use standard [http://en.wikipedia.org/wiki/State_function thermodynamic parameters] of gas, a.k.a state variables of a gas, on which are based quantities like the enthalpy: temperature, volume and pressure. We will also use the velocity because the actual work given by a turbofan engine is related to the mass flow rate of the gas expelled by the engine, which relates to velocity of this gas and the state variables. ===Gas variables: temperature, pressure, velocity=== ===Nomenclature=== Click on the following image to see it larger. All main terms are depicted. [[File:Blade_nomenclature.png|360px|center]] ===Fan design=== The main goal of the fan is to increase the mass flow rate, by increasing the velocity. The mass flow is related to the area of the fan blades and the angular speed of the fan. To increase the velocity, the simplest way is to reduce volume, in other words create a nozzle. The fan duct will thus have to act as a compressor on the aft-end. On the front-end, it is generally designed as an expander, to increase the pressure, allowing more efficient work on the air flow. ===Compressor design=== ''Main article: [[Turbofan:Compressor]]'' The main goal of the compressor is to increase pressure. A compressor stage is composed of a blade-mounted rotor and a vane-mounted stator. The shape of the blade is an airfoil and as the gas flows through the rotor and the stator it gains static pressure, since the blades form a expander. Speed is however gained in the rotor section, because of the high rotational speed of the blades. Stator vanes remove the resulting swirl, and converts the velocity (dynamic pressure) to static pressure, thus increasing again the pressure. The volume occupied by the gas can consequently drop as the pressure increases, intake area will be greater than compressor discharge area. Due to friction and pressure rise, temperature will be increased too. [[File:Compressor_blades.png|360px|center]] ===Turbine design=== The main goal of the turbine is to extract energy from the hot and fast gas discharged by the combustion into mechanical (rotational) work. Turbine vanes generally reduce the discharge area, reducing temperature and increasing speed of the flow then impacting the tubine blades. Pressure and temperature may remain constant ([http://en.wikipedia.org/wiki/Turbine#Theory_of_operation impulse turbine]) or not ([http://en.wikipedia.org/wiki/Turbine#Theory_of_operation reaction turbine]) through the turbine rotor. Recent engines feature a reaction turbine. Sectional area decreasing over the turbine blade help energy to be extracted as rotational work while the flow speed increases. ==Mechanical constraints== Blades on all three parts of a turbofan engine undergo heavy mechanical constraints due to high rpm achieved by the rotor, the high temperature in the turbine section and non-negligible temperature in the end of the compressor section, and the high pressure of the gas on which work is performed. Fan blades are not made of plain metal in real engines. In the eighties, they were made in honeycomb composite sandwich material, they are now made in triangular sandwich structure. For an engine of the size we are targeting, fan blades may be built with plain carbon fiber, like the [http://www.schuebeler-jets.com/index.php?option=com_content&task=view&id=102&Itemid=171 Schuebeler HST] high quality R/C engine. Compressor blades are made of titanium alloys, providing high strength and rigidity at these temperatures. Turbine blades are made of nickel alloys, better sustaining the high temperature, and still at higher strength than steel. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing], as a [http://en.wikipedia.org/wiki/Hot_working hot working] process, is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing hot isostatic pressing] possibly too, as explained on the ''How are made turbine blades'' [http://www.youtube.com/watch?v=vN3_Wkyl5PQ video]. I believe that a hot forging press can be done cheaply considering the small size of our blades. For the main fan, it thus may not be used. [[File:Blade.jpg|600px|center]] ''The above picture lacks the airfoil section of the blade, because the CAD software we use doesn't currently support it.'' Work-hardening and heat-resistance processing have to be studied, and depend on the crafting method used in the first place. The high-pressure turbine blades have to face very high temperature and pressure. On real engines, they are made of titanium and nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. Since the required lifetime is lower in our case, we may achieve a working engine with cheaper metals, like steel or nickel-rich alloys for the turbine blades. Historically, the temperature that turbine blades could sustain greatly improved over time, as well as engine efficiency. We can take these two factors and design a reasonably short-lived (6hrs, 3 missions?) and medium efficiency engine. Material for compressor blade and fan blade can probably be aluminum or aluminum alloys like 2024 (dural) or even 7075. The blade insert will have to be milled at some point, unless if it is casted or pressed. Sonic speed at the tip of a 160mm fan is achieved at around 40000 rpm, for 120mm a fan, required rpm are 50000. Strength of the blade insert and stage's hub should be calculated from this speed and expected blade weights, to verify the capability of materials before fixing engine characteristics. [[Category:Turbofan|Blades]] e5934e9f8970f7c1258f4e255387cd6436662d7d 259 256 2011-11-27T03:52:47Z Vincent 1 /* Nomenclature */ wikipedia nomenclature wikitext text/x-wiki =Blade design and manufacturing= This page explains how blades should be designed for efficiency, and how can a simple and low-cost manufacturing be relevant. ==Blade fixation== The root/fixation/insert of the blade is discussed on the related subsystem design page: [[Turbofan:Compressor|compressor]], [[Turbofan:Turbine|turbine]] or [[Turbofan:Fan|fan]]. In summary, real engines use a fir tree shape to hold the blade centrifugally, while keeping them free axially. They are then generally axially fixed using a locking screw. This design allows for easy replacement of damaged blades, but is quite complex to manufacture, and requires lots of parts. Since our engine will be smaller, using fixing screws is likely to be not feasible, and a single-spool design will make blades subject to less centrifugal stress. Simpler blade fixation mechanism thus had to be studied. ==Blade designs for efficient air flow== Blades have to be carefully designed, because the overall efficiency of the engine largely depends on their design. '''Stages.''' They are not only made of blades on the rotor, but also blades on the stator, generally called ''vanes''. A '''stage''' is then a pair of a rotor stage and a stator stage, n this order for compressor stages, and in the reverse order for turbine stages. Vanes prevent a rotating air flow to form inside the engine (swirl), driven by the action of the rotor blades. Stator vanes redirect the airflow in the more appropriate direction for the next rotor stage. They increase the energy of the gas ([http://en.wikipedia.org/wiki/Enthalpy enthalpy]) by removing the swirling effect that impairs it. '''Blade shape.''' Most basic design of a fan has flat-shaped blades. Twisted blades with a flat section are an improvement, taking into account the difference in apparent airflow velocity and torque all along the blade. Next step is to have a non-flat section, but an airfoil section. This has been [http://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines proved] to be the only design meeting efficiency requirements of the turbine engines, in 1926 by [http://en.wikipedia.org/wiki/Axial_compressor#Development Alan A. Griffith]. Finally, modern engines are designed with curved edges for the fan, for optimal known efficiency as well as for noise reduction. '''Rotor/stator gaps.''' Highest efficiency is reached in turbofans when gaps are reduced between blades and the stator, or between the rotor and stator blades. As always, good efficiency means good high precision and higher costs. Anyway, the precision of blades will have to be very good if we don't want the engine to dislocate when it reaches the high rotations-per-minute achieved. A high reproducibility is required and partially automated manufacturing allows it. Design of stages is linked to the energy the blades have to give (compressor and fan) or take (turbine) to the air flow. To better define and understand that energy, we will use standard [http://en.wikipedia.org/wiki/State_function thermodynamic parameters] of gas, a.k.a state variables of a gas, on which are based quantities like the enthalpy: temperature, volume and pressure. We will also use the velocity because the actual work given by a turbofan engine is related to the mass flow rate of the gas expelled by the engine, which relates to velocity of this gas and the state variables. ===Gas variables: temperature, pressure, velocity=== ===Nomenclature=== Click on the following image to see it larger. All main terms are depicted. [[File:Blade_nomenclature.png|360px|center]] More information can be found on Wikipedia's page for [http://en.wikipedia.org/wiki/Airfoil Airfoil]. ===Fan design=== The main goal of the fan is to increase the mass flow rate, by increasing the velocity. The mass flow is related to the area of the fan blades and the angular speed of the fan. To increase the velocity, the simplest way is to reduce volume, in other words create a nozzle. The fan duct will thus have to act as a compressor on the aft-end. On the front-end, it is generally designed as an expander, to increase the pressure, allowing more efficient work on the air flow. ===Compressor design=== ''Main article: [[Turbofan:Compressor]]'' The main goal of the compressor is to increase pressure. A compressor stage is composed of a blade-mounted rotor and a vane-mounted stator. The shape of the blade is an airfoil and as the gas flows through the rotor and the stator it gains static pressure, since the blades form a expander. Speed is however gained in the rotor section, because of the high rotational speed of the blades. Stator vanes remove the resulting swirl, and converts the velocity (dynamic pressure) to static pressure, thus increasing again the pressure. The volume occupied by the gas can consequently drop as the pressure increases, intake area will be greater than compressor discharge area. Due to friction and pressure rise, temperature will be increased too. [[File:Compressor_blades.png|360px|center]] ===Turbine design=== The main goal of the turbine is to extract energy from the hot and fast gas discharged by the combustion into mechanical (rotational) work. Turbine vanes generally reduce the discharge area, reducing temperature and increasing speed of the flow then impacting the tubine blades. Pressure and temperature may remain constant ([http://en.wikipedia.org/wiki/Turbine#Theory_of_operation impulse turbine]) or not ([http://en.wikipedia.org/wiki/Turbine#Theory_of_operation reaction turbine]) through the turbine rotor. Recent engines feature a reaction turbine. Sectional area decreasing over the turbine blade help energy to be extracted as rotational work while the flow speed increases. ==Mechanical constraints== Blades on all three parts of a turbofan engine undergo heavy mechanical constraints due to high rpm achieved by the rotor, the high temperature in the turbine section and non-negligible temperature in the end of the compressor section, and the high pressure of the gas on which work is performed. Fan blades are not made of plain metal in real engines. In the eighties, they were made in honeycomb composite sandwich material, they are now made in triangular sandwich structure. For an engine of the size we are targeting, fan blades may be built with plain carbon fiber, like the [http://www.schuebeler-jets.com/index.php?option=com_content&task=view&id=102&Itemid=171 Schuebeler HST] high quality R/C engine. Compressor blades are made of titanium alloys, providing high strength and rigidity at these temperatures. Turbine blades are made of nickel alloys, better sustaining the high temperature, and still at higher strength than steel. ==Manufacturing propositions== [http://en.wikipedia.org/wiki/Forging_press Hot pressing], as a [http://en.wikipedia.org/wiki/Hot_working hot working] process, is used to manufacture real-engines' blades, and [http://en.wikipedia.org/wiki/Hot_isostatic_pressing hot isostatic pressing] possibly too, as explained on the ''How are made turbine blades'' [http://www.youtube.com/watch?v=vN3_Wkyl5PQ video]. I believe that a hot forging press can be done cheaply considering the small size of our blades. For the main fan, it thus may not be used. [[File:Blade.jpg|600px|center]] ''The above picture lacks the airfoil section of the blade, because the CAD software we use doesn't currently support it.'' Work-hardening and heat-resistance processing have to be studied, and depend on the crafting method used in the first place. The high-pressure turbine blades have to face very high temperature and pressure. On real engines, they are made of titanium and nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys]. Since the required lifetime is lower in our case, we may achieve a working engine with cheaper metals, like steel or nickel-rich alloys for the turbine blades. Historically, the temperature that turbine blades could sustain greatly improved over time, as well as engine efficiency. We can take these two factors and design a reasonably short-lived (6hrs, 3 missions?) and medium efficiency engine. Material for compressor blade and fan blade can probably be aluminum or aluminum alloys like 2024 (dural) or even 7075. The blade insert will have to be milled at some point, unless if it is casted or pressed. Sonic speed at the tip of a 160mm fan is achieved at around 40000 rpm, for 120mm a fan, required rpm are 50000. Strength of the blade insert and stage's hub should be calculated from this speed and expected blade weights, to verify the capability of materials before fixing engine characteristics. [[Category:Turbofan|Blades]] efd3c785d8d25d7eb609c3bfee5f438b4b7f6012 6 54 255 2011-11-27T01:39:33Z Vincent 1 Velocity triangles for axial compressor blades. Source is the chapter "2.0 Axial-Flow Compressors" from M. P. Broyce. wikitext text/x-wiki Velocity triangles for axial compressor blades. Source is the chapter "2.0 Axial-Flow Compressors" from M. P. Broyce. b54178e963816dbaa95df0ef19ac7449d920b2d4 0 1 261 248 2011-12-05T01:14:21Z Vincent 1 /* The aircraft */ adding text on mig25 and fixes of other blocks wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2012. Anyway, if you find this project interesting, you can still [[Join|join]] and [[Guidelines|participate]]! Maybe with several people we can still make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==Current status of the project / News== Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. Since the main constraint is to have low costs, we'll have to build the plane first, including the turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus focus on the turbofan research and design. The first step is to have the engine [[Build_a_cheap_turbofan|designed]] and built. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are currently being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. '''''(updated November 2011)''''' Information available on this site is sometimes quite outdated, and may reflect some weak or false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design it the fighter jet, for example the [http://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. Can electricity energy be considered for that kind of mission? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, a major issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive and their use is restricted to missiles or UAVs. ===Staging and recovery=== Separation from the rocket is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a complicated autonomous return-to-runway and landing system. ===Guidance=== A satellite navigation system can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 8af1a1ade0d142a4d22d8d200169f27251f9e6a4 263 261 2011-12-09T00:56:38Z Vincent 1 /* The aircraft */ adding link to new page flight at high altitude wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2012. Anyway, if you find this project interesting, you can still [[Join|join]] and [[Guidelines|participate]]! Maybe with several people we can still make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==Current status of the project / News== Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. Since the main constraint is to have low costs, we'll have to build the plane first, including the turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus focus on the turbofan research and design. The first step is to have the engine [[Build_a_cheap_turbofan|designed]] and built. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are currently being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. '''''(updated November 2011)''''' Information available on this site is sometimes quite outdated, and may reflect some weak or false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design it the fighter jet, for example the [http://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. Can electricity energy be considered for that kind of mission? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, a major issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive and their use is restricted to missiles or UAVs. ===Staging and recovery=== Separation from the rocket is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a complicated autonomous return-to-runway and landing system. ===Guidance=== A satellite navigation system can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. eb992c3815ba2aaa189a4d4e2c35de03a70ca66c 269 263 2012-02-09T00:59:36Z Vincent 1 /* Current status of the project / News */ wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2012. Anyway, if you find this project interesting, you can still [[Join|join]] and [[Guidelines|participate]]! Maybe with several people we can still make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==News / current status of the project== Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. Since the main constraint is to have low costs, we'll have to build the plane first, including the turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan research and design. The first step is to have the engine [[Build_a_cheap_turbofan|designed]] and built. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. '''''February 2012 update:''''' Man power will rise from 5 (men.hours)/week to 40 in the next month, the project will have some kind of lift. Study of aerodynamics is under way, and we hope that a preliminary design of the turbofan engine can be completed for the end of March. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may reflect some weak or false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design it the fighter jet, for example the [http://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. Can electricity energy be considered for that kind of mission? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, a major issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive and their use is restricted to missiles or UAVs. ===Staging and recovery=== Separation from the rocket is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a complicated autonomous return-to-runway and landing system. ===Guidance=== A satellite navigation system can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. a39d87970ea238583a7c70e085e2ed0986f31226 280 269 2012-04-03T14:55:31Z Vincent 1 /* News / current status of the project */ wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2012. Anyway, if you find this project interesting, you can still [[Join|join]] and [[Guidelines|participate]]! Maybe with several people we can still make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==News / current status of the project== Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. Since the main constraint is to have low costs, we'll have to build the plane first, including the turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan research and design. The first step is to have the engine [[Build_a_cheap_turbofan|designed]] and built. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may reflect some weak or false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [http://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([http://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [http://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design it the fighter jet, for example the [http://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. Can electricity energy be considered for that kind of mission? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, a major issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive and their use is restricted to missiles or UAVs. ===Staging and recovery=== Separation from the rocket is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a complicated autonomous return-to-runway and landing system. ===Guidance=== A satellite navigation system can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [http://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [http://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [http://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 7c7a36b2b7c7c73cc6ab1117c2f2fde84f594cf8 281 280 2012-04-03T16:10:16Z Vincent 1 https for wikipedia wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2012. Anyway, if you find this project interesting, you can still [[Join|join]] and [[Guidelines|participate]]! Maybe with several people we can still make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==News / current status of the project== Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. Since the main constraint is to have low costs, we'll have to build the plane first, including the turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan research and design. The first step is to have the engine [[Build_a_cheap_turbofan|designed]] and built. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may reflect some weak or false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design it the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. Can electricity energy be considered for that kind of mission? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, a major issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive and their use is restricted to missiles or UAVs. ===Staging and recovery=== Separation from the rocket is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a complicated autonomous return-to-runway and landing system. ===Guidance=== A satellite navigation system can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 15949d6640bf9d7d068912788b89dfc55cc6b6eb 291 281 2012-05-11T22:53:17Z Vincent 1 /* News / current status of the project */ may news wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2012. Anyway, if you find this project interesting, you can still [[Join|join]] and [[Guidelines|participate]]! Maybe with several people we can still make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==News / current status of the project== Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. Since the main constraint is to have low costs, we'll have to build the plane first, including the turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan research and design. The first step is to have the engine [[Build_a_cheap_turbofan|designed]] and built. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may reflect some weak or false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design it the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. Can electricity energy be considered for that kind of mission? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, a major issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive and their use is restricted to missiles or UAVs. ===Staging and recovery=== Separation from the rocket is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a complicated autonomous return-to-runway and landing system. ===Guidance=== A satellite navigation system can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit. I believe that simple cameras can be used on the rocket to determine position of the sun and the Earth's horizon. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used too. Anyway, if sensors are available, actuators are different story. I see only two possibilities, as fins won't have any impact in the vacuum of space: the rocket engine has to be directionally controllable or control jets must be used to control the attitude of the rocket, as does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but seem mandatory. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 17bf3a5cf15e8853978142a00337fa7305faea62 292 291 2012-05-11T23:12:50Z Vincent 1 /* Trajectory */ more details on attitude control and actuators wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise to actually build it in time before the deadline of the project in september 2012. Anyway, if you find this project interesting, you can still [[Join|join]] and [[Guidelines|participate]]! Maybe with several people we can still make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==News / current status of the project== Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. Since the main constraint is to have low costs, we'll have to build the plane first, including the turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan research and design. The first step is to have the engine [[Build_a_cheap_turbofan|designed]] and built. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may reflect some weak or false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design it the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. Can electricity energy be considered for that kind of mission? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, a major issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive and their use is restricted to missiles or UAVs. ===Staging and recovery=== Separation from the rocket is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a complicated autonomous return-to-runway and landing system. ===Guidance=== A satellite navigation system can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. I believe that mere cameras can be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. Anyway, if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: 1) the rocket engine has to be directionally controllable (generally using hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]), or 2) control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) must be used to control the attitude of the rocket, as partially does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but are mandatory in our case. This is one of the big differences between sub-orbital and orbital space flight. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. e81c1f71c0abc1b0469c4d6412ae412efc1ad580 294 292 2012-05-27T15:37:26Z Vincent 1 remade the news and project description sections, adding the boeing news wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, you can [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide a low cost orbital launch system for small size satellites, probably with a mass lower than 1kg. Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. The rocket, as in any other orbital launch system, would make it to orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. ==News== '''''May 21, 2012: ''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may reflect some weak or false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design it the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. Can electricity energy be considered for that kind of mission? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, a major issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive and their use is restricted to missiles or UAVs. ===Staging and recovery=== Separation from the rocket is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a complicated autonomous return-to-runway and landing system. ===Guidance=== A satellite navigation system can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. I believe that mere cameras can be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. Anyway, if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: 1) the rocket engine has to be directionally controllable (generally using hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]), or 2) control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) must be used to control the attitude of the rocket, as partially does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but are mandatory in our case. This is one of the big differences between sub-orbital and orbital space flight. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 1fd34241e234010539d389ceeae4826c29369c85 0 49 262 249 2011-12-05T01:57:54Z Vincent 1 /* Join the project! */ otaski link and fixes wikitext text/x-wiki =Join the project!= One of the missions of this project is to provide easy access to complicated information and hard to find. Technologies are studied and compared, and results are clearly explained in order to explain why we have to chose one solution instead of another. Innovative solutions are also proposed and evaluated. All resources are made available freely, in accordance with [http://otaski.org OTASKI], check the licence here or there. ==Why join?== Lots of technologies and scientific fields of research are covered by aerospace activity. Help is welcome, because it takes too much years to understand deeply all the details of each field. If you can provide information, resources or if you can orientate the choices made for the prototypes, [[Founder|I]]'d be glad to have you the project. ==How to join?== [[Special:RequestAccount|Creating an account]] on the wiki allows you to participate to the project. It has to be accepted by administrators before you can use it. Account requests have to be completed with a minimum of information about what you would like to do in the project or in what domain you have some expertise. Thank you! 3264bdc8fda3e1cb535679b26666a5308d005f0a 0 55 264 2011-12-09T00:57:43Z Vincent 1 creating page wikitext text/x-wiki =Flight at high altitude= Some information is summarized in the main page already, in the [[Main_Page#The_aircraft|aircraft section]]. ==Approaches overview== ==High engine power at low air density== ==High lift at low air density== 089fe88dea1bd8f875d9ff30d37c042358b85632 266 264 2011-12-09T01:13:25Z Vincent 1 /* Flight at high altitude */ some links and first interrogations wikitext text/x-wiki =Flight at high altitude= Some information is summarized in the main page already, in the [[Main_Page#The_aircraft|aircraft section]]. ==Gas properties and altitude== Density is used to calculate [http://en.wikipedia.org/wiki/Lift_(force) lift] of an wing and [http://en.wikipedia.org/wiki/Thrust thrust] of an engine amongst other. We need to know what air densities will be faced in order to design everything. Atmospheric density depends on pressure. [http://www.respirometry.org/index.php/look-up-tables/37-lookup-tables-cat/58-barometric-pressure-vs-altitude- This table] gives atmospheric pressure and temperature depending on altitude. We can see that a tenth of ground atmospheric pressure is met at around 16km altitude, and a hundredth of it at around 31km altitude. ==Approaches overview== ==High engine power at low air density== '''Is it possible to have a low total pressure ratio engine operating at subsonic inlet speeds?''' The MiG 25 has supersonic inlet, which allows him to have a significant pressure increase before the compressor actually gives energy to the flow. ==High lift at low air density== 1e140b9ad8f37620b94abbb12b82f381b689332d 267 266 2011-12-19T02:08:58Z Vincent 1 more on air density wikitext text/x-wiki =Flight at high altitude= Some information is summarized in the main page already, in the [[Main_Page#The_aircraft|aircraft section]]. ==Gas properties and altitude== Density is used to calculate [http://en.wikipedia.org/wiki/Lift_(force) lift] of an wing and [http://en.wikipedia.org/wiki/Thrust thrust] of an engine amongst other. We need to know what air densities will be faced in order to design everything. Air density depends on pressure. [http://www.respirometry.org/index.php/look-up-tables/37-lookup-tables-cat/58-barometric-pressure-vs-altitude- This table] gives atmospheric pressure and temperature depending on altitude. We can see that a tenth of ground atmospheric pressure is met at around 16km altitude, and a hundredth of it at around 31km altitude. Air density in the atmosphere is also related to the ratio of air vapour in it, as indicated on [http://wahiduddin.net/calc/density_altitude.htm this page]. The page also contains lots of formulas and calculators, most importantly the ''air density calculator'' that we'll use right below. Water vapour however, is much more rare when temperature goes down, as it does in the higher troposphere or low to mid stratosphere that we're aiming. The calculator gives us, with temperature and pressure values taken from the table mentioned above, values for density of: * 1.214 kg/m^3 at sea level * 0.1877 kg/m^3 at 15km altitude * 0.0441 kg/m^3 at 25km altitude * 0.017 kg/m^3 at 30km altitude ==Approaches overview== ==High engine power at low air density== '''Is it possible to have a low total pressure ratio engine operating at subsonic inlet speeds and low air density?''' The MiG 25 has supersonic inlet, which allows him to have a significant pressure increase before the compressor actually gives energy to the flow. A subsonic input air flow in the high-altitude conditions is likely to not provide enough oxygen for the combustion to maintain by itself, or a too poor mass flow rate to the turbine. ==High lift at low air density== High engine power in low air density generally means supersonic flight, or at least, high flight speeds, which in return increase the lift of the aircraft or decrease its wingspan. Our next step is to make some calculations of the required winged area for subsonic low-density air travel, and assess the feasibility of our air launch to orbit project. e28d5abbfd75dce1ad6bc8ea125aaf8a8d8ef881 0 11 265 232 2011-12-09T01:00:06Z Vincent 1 /* Turbofan design procedure */ links to work on the three first design steps wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets or other engines. They are however very difficult to design and manufacture and are thus very expensive. On this page, we will explore how costs can be reduced while still having adequate capabilities for high altitude flight. ==General principles== Lots of information are available on [http://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that puts energy into a gas, this energy is extracted by a turbine, and the turbine drives both the fan that provides thrust and the compression stage that feeds the combustion with oxygen. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake, and so on. The fan provides thrust by creating a massive air flow, and the engine's core also creates thrust by evacuating the high-speed hot combustion gas. In commercial turbofan engines, the fan is generally responsible for 90% of the overall thrust. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [http://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost, has the fan mounted on it. One stage engines exist and are less complicated and expensive to build, but are much less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan, but this is not an easy task due to this very speed. Commercial engines featuring a gearbox for the turbofan's fan are expected to reach market in 2012. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan, but are not yet optimal. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seeing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down with the cost. ===Shaped core or shaped shaft?=== An obvious but important optimization to reduce cost and complexity of manufacturing is to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to design and produce, with a simple [http://en.wikipedia.org/wiki/Lathe lathe] ([http://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [http://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates, for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a shaft turned in raw metal won't change much on its final mass. Moreover, we may use a turbine-level mechanism embedded in the stator to try to cool it, which would make it hollow. The main mechanical issues are probably how to properly fix the blades on rotor and stator, how to fix the rotor on the stator with little gap, and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compressor blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [http://en.wikipedia.org/wiki/Superalloys superalloys] or are ceramic-coated. It's the inability of blades to withstand heat and work that limit the power of the engine. Indeed, around 70% of the gas provided by the compressor is used only for chamber and turbine cooling, instead of using it to burn more fuel and create more thrust. The compressor and the turbine are not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow driven by the action of rotor blades to form inside the engine, which would decrease the energy of the gas. Stator blades or vanes redirect the airflow on the next stage in the more efficient direction. Highest efficiencies are reached in turbofans when gaps are reduced between rotor blades' tip and the stator, as well as between the stator blades' tip and the rotor. As always, good efficiency means high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. The shape of the blade and the parameters of their cascade also affects the efficiency. A small 5 stage supersonic compressor providing the same pressure rise than a 15 stage subsonic compressor is less efficient, but it may be compensated by the higher thrust-to-weight ratio. Blade geometric design is also very complicated. First turbine engines had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1926, [http://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines Alan A. Griffith] proved that if blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to deserve being built. Airfoils for blade designs allow compressor stages to better increase the static pressure since they create an expander, an increasing area for the air flow to pass through. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C so it will be way off-limits, and even below that, it will deform before melting. And should we mention that blade deformation at high-centrifugal loads, caused by the high rpm, is a very good way to blow the engine off? Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the vanes the turbine, which are, on actual engines, made of a ceramic-coated high-temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge or bleed circulates, enters the blades, and evacuates through small holes in the blades (convective cooling and film cooling). For the rotor blades, the same principle is used, but with compressor air passing inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air or a high speed electric engine for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air and compressor stall, re-ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude. The rotor speed information and altimeter may be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and removed axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we aim. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, a simple square-section root is probably enough. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blade-roots are inserted and fixed. Creating a perpendicular blade root is already a challenge. Rotor blades would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting should be considered. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Air bled from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose [[Turbofan:Alternative Designs|alternative turbofan designs]] based on a mixed-flow turbine. Several pages have been created in the [[:Category:Turbofan|Turbofan category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three to five stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while sustaining the combustion. * [[Turbofan:Bearings|Bearings and cooling]]: high speed rotations require adapted bearings and cooling, which may be reused for rotor and even turbine cooling. ===Turbofan design procedure=== # evaluate required thrust (from aircraft mass and lift, but also [[Flight_at_high_altitude|flight characteristics]]) # calculate required mass flow rate for the fan (thust is [http://en.wikipedia.org/wiki/Thrust calculated] from MFR and flow speed) # fix bypass ratio and fan diameter and rpm, thus giving core diameter (BPR may be [[Turbofan:Alternative_Designs#Full_transonic_engine_design_in_a_single_spool_with_2.1_BPR|fixed by design]]) # calculate required power to drive the fan alone # evaluate a gross compressor driving power (refined later)<br />&nbsp; # calculate total power that has to be drawn from the turbine (fan + compressor + losses) # calculate mass flow rate for the combustion alone # calculate mass flow rate for cooling chamber and turbine ## calculate mass flow rate for cooling chamber ## evaluate mass flow rate for cooling turbine to add to the latter ## calculate entropy and fluid parameters at combustor discharge (speed, temperature) ## calculate temperature of turbine vanes and blade and check if it is acceptable ## iterate on item 8.1 until temperature is unacceptable # calculate the number of turbine blades and stages required for this power # calculate compressor discharge pressure and pressure ratio # calculate how many compressor stages are required depending on sonic or supersonic blade design and fix design # calculate compressor driving power # iterate on item 6 until total power varies # design blades for all calculated parameters and re-run at item 6, total power may have changed [[Category:Turbofan]] c7dbcc641b54c2f3f31356efccb4cc11822a12c2 0 16 268 237 2011-12-31T18:05:18Z Vincent 1 new links on micro rocketry and aerodynamics wikitext text/x-wiki =Resources= This page gathers available documentation on the numerous subjects linked the project, including rocket science, turbofans, aerodynamics, amateur rocketry and SSTO. ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/l amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * An important nasaspaceflight [http://forum.nasaspaceflight.com/index.php?topic=2847.0 forum thread] where ideas on micro-rocketry to orbit are discussed and shared. * A kind of spin-off of the above thread is the [http://orbitalaspirations.blogspot.com/ orbital aspirations weblog]. It was recently created by Ed LeBouthillier and is already filled with lots of information on scaling down rocket equations to micro-rocketry, SSTO, reference papers, news of the domain and so on. * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] by Klaus Hünecke, 1997. * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ===Other Aerodynamics=== * [http://www.desktop.aero/appliedaero/preface/welcome.html Applied Aerodynamics: A Digital Textbook] by Desktop Aeronautics, Inc. 2007. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. 0d46f550e8c38302a692082b7ddbc061372874a1 271 268 2012-03-15T00:13:57Z Vincent 1 adding internal resource link wikitext text/x-wiki =Resources= This page gathers available documentation on the numerous subjects linked the project, including rocket science, turbofans, aerodynamics, amateur rocketry and SSTO. An [[Aero formulas|internal page]] is being created containing useful formulas related to aerodynamics. ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/l amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * An important nasaspaceflight [http://forum.nasaspaceflight.com/index.php?topic=2847.0 forum thread] where ideas on micro-rocketry to orbit are discussed and shared. * A kind of spin-off of the above thread is the [http://orbitalaspirations.blogspot.com/ orbital aspirations weblog]. It was recently created by Ed LeBouthillier and is already filled with lots of information on scaling down rocket equations to micro-rocketry, SSTO, reference papers, news of the domain and so on. * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] by Klaus Hünecke, 1997. * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ===Other Aerodynamics=== * [http://www.desktop.aero/appliedaero/preface/welcome.html Applied Aerodynamics: A Digital Textbook] by Desktop Aeronautics, Inc. 2007. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. bb151b619273537f128c9a2443633971fbd3399e 0 56 272 2012-03-15T00:50:35Z Vincent 1 init tables with ideal gas formula wikitext text/x-wiki =Resources on physics related to aerodynamics= The [http://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Value (if any) !Unit (SI) |- | | | |- | P | Pressure | | Pa (pascal) |- | T | Temperature | | K (kelvin) |- | V | Volume | | l (litre) |- | n | quantity of matter | | mole |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value (if any) !Unit (SI) |- | R | [http://en.wikipedia.org/wiki/Gas_constant perfect gas constant] | 8.3144621 | J K−1 mol−1 |} 68b892de563b2018051aa1dde42f1484ab610a13 273 272 2012-03-15T00:58:35Z Vincent 1 Super scripts and fixes wikitext text/x-wiki =Resources on physics related to aerodynamics= The [http://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | | |- | P | Pressure | Pa (pascal) |- | T | Temperature | K (kelvin) |- | V | Volume | m³ (cubic meter) |- | n | Quantity of matter | mol (mole) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value (if any) !Unit (SI) |- | R | [http://en.wikipedia.org/wiki/Gas_constant perfect gas constant] | 8.3144621 | J K⁻¹ mol⁻¹ |} 3f0dec8bd22ad6bdaa5274ce928ad7eeb6a78d5e 274 273 2012-03-15T02:19:53Z Vincent 1 more variables and constants, images wikitext text/x-wiki =Resources on physics related to aerodynamics= The [http://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | H | Enthalpy | J (Joule) |- | style="background:white; color:black"| {{SERVER}}/images/formulas_mirror/heat_vap.png or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 |- | T | Temperature | K (Kelvin) |- | V | Volume | m³ (cubic meter) |- | n | Quantity of matter | mol (mole) |- | p | Pressure | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). |- | | | |} fa717ac73f751dbecc947842f119fe9cd9fac974 275 274 2012-03-15T02:36:50Z Vincent 1 clausius wikitext text/x-wiki =Resources on physics related to aerodynamics= The [http://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | H | Enthalpy | J (Joule) |- | style="background:white; color:black"| {{SERVER}}/images/formulas_mirror/heat_vap.png or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 |- | T | Temperature | K (Kelvin) |- | V | Volume | m³ (cubic meter) |- | n | Quantity of matter | mol (mole) |- | p | Pressure | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures. |} a4e45fca0d73e234abf896f978f7c208e6be4660 276 275 2012-03-15T02:54:08Z Vincent 1 surface tension, viscosity wikitext text/x-wiki =Resources on physics related to aerodynamics= The [http://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | H | Enthalpy | J (Joule) |- | style="background:white; color:black"| {{SERVER}}/images/formulas_mirror/heat_vap.png or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | T | Temperature | K (Kelvin) |- | V | Volume | m³ (cubic meter) |- | n | Quantity of matter | mol (mole) |- | p | Pressure | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |} 5adbf3ef412d5f0ed0a9dcec49d864bea92a00b7 277 276 2012-03-15T03:02:31Z Vincent 1 /* List of variables */ enthalpy link wikitext text/x-wiki =Resources on physics related to aerodynamics= The [http://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: energy of a thermodynamic system. | J (Joule) |- | style="background:white; color:black"| {{SERVER}}/images/formulas_mirror/heat_vap.png or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | T | Temperature | K (Kelvin) |- | V | Volume | m³ (cubic meter) |- | n | Quantity of matter | mol (mole) |- | p | Pressure | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |} bd687aadc80eeb899386ed1f1be44383bdd0f453 278 277 2012-03-29T01:09:22Z Vincent 1 Heat equations wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: energy of a thermodynamic system. | J (Joule) |- | style="background:white; color:black"| {{SERVER}}/images/formulas_mirror/heat_vap.png or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature | K (Kelvin) |- | V | Volume | m³ (cubic meter) |- | n | Quantity of matter | mol (mole) |- | p | Pressure | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/Qdefinition.png |Definition of [https://en.wikipedia.org/wiki/Heat#Path-independent_examples_for_an_ideal_gas Heat] for an ideal gas. |The heat required to change the temperature of a system from an initial temperature T₀, to a final temperature, T_f. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/QeqmL.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. |- |} e22a8093d3a6a694bd43c3a23eb28183b9a98151 279 278 2012-03-29T02:21:01Z Vincent 1 First law of thermodynamics wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: total energy of a thermodynamic system. | J (Joule) |- | style="background:white; color:black"| {{SERVER}}/images/formulas_mirror/heat_vap.png or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature | K (Kelvin) |- | U | [https://en.wikipedia.org/wiki/Internal_energy Internal energy] of a system (see first law of Thermodynamics below) | J (Joule) |- | V | Volume | m³ (cubic meter) |- | W | [https://en.wikipedia.org/wiki/Work_(thermodynamics) Work]: mechanical constraints on the system. | J (Joule) |- | n | Quantity of matter | mol (mole) |- | p | Pressure | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/Qdefinition.png |Definition of [https://en.wikipedia.org/wiki/Heat#Path-independent_examples_for_an_ideal_gas Heat] for an ideal gas. |The heat required to change the temperature of a system from an initial temperature T₀, to a final temperature, T_f. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/QeqmL.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. Delta H equals Q only when pressure is constant (isobaric). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dUeqdQmindW.png |[https://en.wikipedia.org/wiki/First_law_of_thermodynamics First law of Thermodynamics] |Variations of internal energy of a system between two states is the sum of the received heat and work (minus the ''given'' work). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/enthalpy.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] |Total amount of energy of a system, defined as the sum of the internal energy U and pressure * volume. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/workExpand.png |Work of gas expansion. |Work done by expanding an ideal gas. |} d2cde6976c46a8e5c4188a5d5b5eb3cae9732f94 282 279 2012-04-10T02:28:39Z Vincent 1 basic entropy wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: total energy of a thermodynamic system. | J (Joule) |- | style="background:white; color:black"| {{SERVER}}/images/formulas_mirror/heat_vap.png or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature | K (Kelvin) |- | S | [https://en.wikipedia.org/wiki/Entropy Entropy] | J.K^-1 (Joule per Kelvin) |- | U | [https://en.wikipedia.org/wiki/Internal_energy Internal energy] of a system (see first law of Thermodynamics below) | J (Joule) |- | V | Volume | m³ (cubic meter) |- | W | [https://en.wikipedia.org/wiki/Work_(thermodynamics) Work]: mechanical constraints on the system. | J (Joule) |- | n | Quantity of matter | mol (mole) |- | p | Pressure | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/Qdefinition.png |Definition of [https://en.wikipedia.org/wiki/Heat#Path-independent_examples_for_an_ideal_gas Heat] for an ideal gas. |The heat required to change the temperature of a system from an initial temperature T₀, to a final temperature, T_f. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/QeqmL.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. Delta H equals Q only when pressure is constant (isobaric). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dUeqdQmindW.png |[https://en.wikipedia.org/wiki/First_law_of_thermodynamics First law of Thermodynamics] |Variations of internal energy of a system between two states is the sum of the received heat and work (minus the ''given'' work). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/enthalpy.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] |Total amount of energy of a system, defined as the sum of the internal energy U and pressure * volume. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/workExpand.png |Work of gas expansion. |Work done by expanding an ideal gas. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/entropy_dueqtdsmpdv.png |[https://en.wikipedia.org/wiki/Entropy Entropy] |Internal energy related to entropy variation for a closed system in thermal equilibrium ([https://en.wikipedia.org/wiki/Fundamental_thermodynamic_relation fundamental thermodynamic relation]). |} 50d6bb96e7dcb0b999bb057fae38edb48f79767c 283 282 2012-04-10T02:44:30Z Vincent 1 more entropy and gibbs energy wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: total energy of a thermodynamic system. | J (Joule) |- | style="background:white; color:black"| {{SERVER}}/images/formulas_mirror/heat_vap.png or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature | K (Kelvin) |- | S | [https://en.wikipedia.org/wiki/Entropy Entropy] | J.K^-1 (Joule per Kelvin) |- | U | [https://en.wikipedia.org/wiki/Internal_energy Internal energy] of a system (see first law of Thermodynamics below) | J (Joule) |- | V | Volume | m³ (cubic meter) |- | W | [https://en.wikipedia.org/wiki/Work_(thermodynamics) Work]: mechanical constraints on the system. | J (Joule) |- | n | Quantity of matter | mol (mole) |- | p | Pressure | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/Qdefinition.png |Definition of [https://en.wikipedia.org/wiki/Heat#Path-independent_examples_for_an_ideal_gas Heat] for an ideal gas. |The heat required to change the temperature of a system from an initial temperature T₀, to a final temperature, T_f. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/QeqmL.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. Delta H equals Q only when pressure is constant (isobaric). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dUeqdQmindW.png |[https://en.wikipedia.org/wiki/First_law_of_thermodynamics First law of Thermodynamics] |Variations of internal energy of a system between two states is the sum of the received heat and work (minus the ''given'' work). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/enthalpy.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] |Total amount of energy of a system, defined as the sum of the internal energy U and pressure * volume. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/workExpand.png |Work of gas expansion. |Work done by expanding an ideal gas. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/entropy_dueqtdsmpdv.png |[https://en.wikipedia.org/wiki/Entropy Entropy] |Internal energy related to entropy variation for a closed system in thermal equilibrium ([https://en.wikipedia.org/wiki/Fundamental_thermodynamic_relation fundamental thermodynamic relation]). |- |ΔS_universe = ΔS_surroundings + ΔS_system |Entropy variation as a whole. |Entropy variation of a system is generally compensated by the inverse variation of the surroundings not including losses. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/gibbs.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] / Free enthalpy |Useful work obtainable from a system at isobaric and isothermal conditions. |} be072d35fe5add440eed1308ccf412b5746c928c 284 283 2012-04-10T02:46:28Z Vincent 1 fix delta H image to text wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: total energy of a thermodynamic system. | J (Joule) |- | ΔH_vap or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature | K (Kelvin) |- | S | [https://en.wikipedia.org/wiki/Entropy Entropy] | J.K^-1 (Joule per Kelvin) |- | U | [https://en.wikipedia.org/wiki/Internal_energy Internal energy] of a system (see first law of Thermodynamics below) | J (Joule) |- | V | Volume | m³ (cubic meter) |- | W | [https://en.wikipedia.org/wiki/Work_(thermodynamics) Work]: mechanical constraints on the system. | J (Joule) |- | n | Quantity of matter | mol (mole) |- | p | Pressure | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/Qdefinition.png |Definition of [https://en.wikipedia.org/wiki/Heat#Path-independent_examples_for_an_ideal_gas Heat] for an ideal gas. |The heat required to change the temperature of a system from an initial temperature T₀, to a final temperature, T_f. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/QeqmL.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. Delta H equals Q only when pressure is constant (isobaric). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dUeqdQmindW.png |[https://en.wikipedia.org/wiki/First_law_of_thermodynamics First law of Thermodynamics] |Variations of internal energy of a system between two states is the sum of the received heat and work (minus the ''given'' work). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/enthalpy.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] |Total amount of energy of a system, defined as the sum of the internal energy U and pressure * volume. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/workExpand.png |Work of gas expansion. |Work done by expanding an ideal gas. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/entropy_dueqtdsmpdv.png |[https://en.wikipedia.org/wiki/Entropy Entropy] |Internal energy related to entropy variation for a closed system in thermal equilibrium ([https://en.wikipedia.org/wiki/Fundamental_thermodynamic_relation fundamental thermodynamic relation]). |- |ΔS_universe = ΔS_surroundings + ΔS_system |Entropy variation as a whole. |Entropy variation of a system is generally compensated by the inverse variation of the surroundings not including losses. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/gibbs.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] / Free enthalpy |Useful work obtainable from a system at isobaric and isothermal conditions. |} 793c1793b4a48a15d8f2ab66a99fc6d46afbb8e0 285 284 2012-04-10T23:29:56Z Vincent 1 more entropy and variables wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | C, C_p, C_V | [https://en.wikipedia.org/wiki/Heat_capacity#Metrology Heat capacity], general, at constant pressure, at constant volume. | J.K^-1 (Joule per Kelvin) |- | G | [https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] | J (Joule) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: total energy of a thermodynamic system. | J (Joule) |- | ΔH_vap or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature | K (Kelvin) |- | S | [https://en.wikipedia.org/wiki/Entropy Entropy] | J.K^-1 (Joule per Kelvin) |- | U | [https://en.wikipedia.org/wiki/Internal_energy Internal energy] of a system (see first law of Thermodynamics below) | J (Joule) |- | V | Volume | m³ (cubic meter) |- | W | [https://en.wikipedia.org/wiki/Work_(thermodynamics) Work]: mechanical constraints on the system. | J (Joule) |- | n | Quantity of matter | mol (mole) |- | p | Pressure | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/QeqmL.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. Delta H equals Q only when pressure is constant (isobaric). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dUeqdQmindW.png |[https://en.wikipedia.org/wiki/First_law_of_thermodynamics First law of Thermodynamics] |Variations of internal energy of a system between two states is the sum of the received heat and work (minus the ''given'' work). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/enthalpy.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] |Total amount of energy of a system, defined as the sum of the internal energy U and pressure * volume. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/workExpand.png |Work of gas expansion. |Work done by expanding an ideal gas. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/entropy_dueqtdsmpdv.png |[https://en.wikipedia.org/wiki/Entropy Entropy] |Internal energy related to entropy variation for a closed system in thermal equilibrium ([https://en.wikipedia.org/wiki/Fundamental_thermodynamic_relation fundamental thermodynamic relation]). |- |ΔS_universe = ΔS_surroundings + ΔS_system |Entropy variation as a whole. |Entropy variation of a system is generally compensated by the inverse variation of the surroundings, not including losses. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/gibbs.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] / Free enthalpy |Useful work obtainable from a system at isobaric and isothermal conditions. Since H is U + pV, it can be replaced in the equation, making G = H - TS. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/deltaG.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] variation. |This derivation is only valid at constant temperature. |} a5c1f3c75f4dc5f8aec008cf3747f79e5c0e40e7 286 285 2012-04-11T00:25:05Z Vincent 1 more thermodynamics wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | C, C_p, C_V | [https://en.wikipedia.org/wiki/Heat_capacity#Metrology Heat capacity], general, at constant pressure, at constant volume. | J.K^-1 (Joule per Kelvin) |- | G | [https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] | J (Joule) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: total energy of a thermodynamic system. | J (Joule) |- | ΔH_vap or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature | K (Kelvin) |- | S | [https://en.wikipedia.org/wiki/Entropy Entropy] | J.K^-1 (Joule per Kelvin) |- | U | [https://en.wikipedia.org/wiki/Internal_energy Internal energy] of a system (see first law of Thermodynamics below) | J (Joule) |- | V | Volume | m³ (cubic meter) |- | W | [https://en.wikipedia.org/wiki/Work_(thermodynamics) Work]: mechanical constraints on the system. | J (Joule) |- | n | Quantity of matter | mol (mole) |- | p | Pressure | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/QeqmL.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. Delta H equals Q only when pressure is constant (isobaric). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dUeqdQmindW.png |[https://en.wikipedia.org/wiki/First_law_of_thermodynamics First law of thermodynamics] |Variations of internal energy of a system between two states is the sum of the received heat and work (minus the ''given'' work). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/enthalpy.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] |Total amount of energy of a system, defined as the sum of the internal energy U of the system and pressure * volume at the boundary of the system and its environment. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/workExpand.png |Work of gas expansion. |Work done by expanding an ideal gas. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/entropy_dueqtdsmpdv.png |[https://en.wikipedia.org/wiki/Entropy Entropy] |Internal energy related to entropy variation for a closed system in thermal equilibrium ([https://en.wikipedia.org/wiki/Fundamental_thermodynamic_relation fundamental thermodynamic relation]). |- |ΔS_universe = ΔS_surroundings + ΔS_system |Entropy variation as a whole. |Entropy variation of a system is generally compensated by the inverse variation of the surroundings, not including losses. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dS.png |[https://en.wikipedia.org/wiki/Second_law_of_thermodynamics Second law of thermodynamics] |A change in the entropy of a system is the infinitesimal transfer of heat to a closed system driving a reversible process, divided by the equilibrium temperature of the system. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/gibbs.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] / Free enthalpy |Useful work obtainable from a system at '''isobaric''' and '''isothermal''' conditions. Since H is U + pV, it can be replaced in the equation, making G = H - TS. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/deltaG.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] variation. |If ΔG < 0, the system's transformation can be spontaneous, if ΔG = 0 the transformation is inversible and the system is in an equilibrium state, if ΔG > 0 it can't be spontaneous. |} 229047151c2aff9b9b0ee170d88605cd5f7312c6 287 286 2012-04-18T02:38:40Z Vincent 1 stagnation and mach number wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] or [https://en.wikipedia.org/wiki/Heat_capacity_ratio Heat capacity ratio] (adiabatic process in thermodynamics) | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | C, C_p, C_V | [https://en.wikipedia.org/wiki/Heat_capacity#Metrology Heat capacity], general, at constant pressure, at constant volume. | J.K^-1 (Joule per Kelvin) |- | G | [https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] | J (Joule) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: total energy of a thermodynamic system. | J (Joule) |- | ΔH_vap or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | M | [https://en.wikipedia.org/wiki/Mach_number Mach number] | no unit |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature. T₀ or T_t is the [https://en.wikipedia.org/wiki/Stagnation_temperature stagnation temperature]. | K (Kelvin) |- | S | [https://en.wikipedia.org/wiki/Entropy Entropy] | J.K^-1 (Joule per Kelvin) |- | U | [https://en.wikipedia.org/wiki/Internal_energy Internal energy] of a system (see first law of Thermodynamics below) | J (Joule) |- | V | Volume | m³ (cubic meter) |- | W | [https://en.wikipedia.org/wiki/Work_(thermodynamics) Work]: mechanical constraints on the system. | J (Joule) |- | a | [https://en.wikipedia.org/wiki/Speed_of_sound Speed of sound] in medium (used to calculate Mach number) | m.s^-1 |- | c | Velocity of a flow in thermodynamics, also noted V; generally noted u in fluid dynamics. | m.s^-1 |- | n | Quantity of matter | mol (mole) |- | p | Pressure. p_t is the [https://en.wikipedia.org/wiki/Stagnation_pressure stagnation pressure]. | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/QeqmL.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. Delta H equals Q only when pressure is constant (isobaric). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dUeqdQmindW.png |[https://en.wikipedia.org/wiki/First_law_of_thermodynamics First law of thermodynamics] |Variations of internal energy of a system between two states is the sum of the received heat and work (minus the ''given'' work). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/enthalpy.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] |Total amount of energy of a system, defined as the sum of the internal energy U of the system and pressure * volume at the boundary of the system and its environment. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/workExpand.png |Work of gas expansion. |Work done by expanding an ideal gas. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/entropy_dueqtdsmpdv.png |[https://en.wikipedia.org/wiki/Entropy Entropy] |Internal energy related to entropy variation for a closed system in thermal equilibrium ([https://en.wikipedia.org/wiki/Fundamental_thermodynamic_relation fundamental thermodynamic relation]). |- |ΔS_universe = ΔS_surroundings + ΔS_system |Entropy variation as a whole. |Entropy variation of a system is generally compensated by the inverse variation of the surroundings, not including losses. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dS.png |[https://en.wikipedia.org/wiki/Second_law_of_thermodynamics Second law of thermodynamics] |A change in the entropy of a system is the infinitesimal transfer of heat to a closed system driving a reversible process, divided by the equilibrium temperature of the system. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/gibbs.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] / Free enthalpy |Useful work obtainable from a system at '''isobaric''' and '''isothermal''' conditions. Since H is U + pV, it can be replaced in the equation, making G = H - TS. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/deltaG.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] variation. |If ΔG < 0, the system's transformation can be spontaneous, if ΔG = 0 the transformation is inversible and the system is in an equilibrium state, if ΔG > 0 it can't be spontaneous. |} 3ed797693f0906e4499a7594c05f0e58ebd385dd 289 287 2012-04-20T01:39:30Z Vincent 1 density wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] or [https://en.wikipedia.org/wiki/Heat_capacity_ratio Heat capacity ratio] (adiabatic process in thermodynamics) | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | ρ (rho) | [https://en.wikipedia.org/wiki/Density Density] | kg.m^-3 (kg per cubic meter) |- | C, C_p, C_V | [https://en.wikipedia.org/wiki/Heat_capacity#Metrology Heat capacity], general, at constant pressure, at constant volume. | J.K^-1 (Joule per Kelvin) |- | G | [https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] | J (Joule) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: total energy of a thermodynamic system. | J (Joule) |- | ΔH_vap or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | M | [https://en.wikipedia.org/wiki/Mach_number Mach number] | no unit |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature. T₀ or T_t is the [https://en.wikipedia.org/wiki/Stagnation_temperature stagnation temperature]. | K (Kelvin) |- | S | [https://en.wikipedia.org/wiki/Entropy Entropy] | J.K^-1 (Joule per Kelvin) |- | U | [https://en.wikipedia.org/wiki/Internal_energy Internal energy] of a system (see first law of Thermodynamics below) | J (Joule) |- | V | Volume | m³ (cubic meter) |- | W | [https://en.wikipedia.org/wiki/Work_(thermodynamics) Work]: mechanical constraints on the system. | J (Joule) |- | a | [https://en.wikipedia.org/wiki/Speed_of_sound Speed of sound] in medium (used to calculate Mach number) | m.s^-1 |- | c | Velocity of a flow in thermodynamics, also noted V; generally noted u in fluid dynamics. | m.s^-1 |- | n | Quantity of matter | mol (mole) |- | p | Pressure. p_t is the [https://en.wikipedia.org/wiki/Stagnation_pressure stagnation pressure]. | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/QeqmL.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. Delta H equals Q only when pressure is constant (isobaric). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dUeqdQmindW.png |[https://en.wikipedia.org/wiki/First_law_of_thermodynamics First law of thermodynamics] |Variations of internal energy of a system between two states is the sum of the received heat and work (minus the ''given'' work). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/enthalpy.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] |Total amount of energy of a system, defined as the sum of the internal energy U of the system and pressure * volume at the boundary of the system and its environment. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/workExpand.png |Work of gas expansion. |Work done by expanding an ideal gas. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/entropy_dueqtdsmpdv.png |[https://en.wikipedia.org/wiki/Entropy Entropy] |Internal energy related to entropy variation for a closed system in thermal equilibrium ([https://en.wikipedia.org/wiki/Fundamental_thermodynamic_relation fundamental thermodynamic relation]). |- |ΔS_universe = ΔS_surroundings + ΔS_system |Entropy variation as a whole. |Entropy variation of a system is generally compensated by the inverse variation of the surroundings, not including losses. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dS.png |[https://en.wikipedia.org/wiki/Second_law_of_thermodynamics Second law of thermodynamics] |A change in the entropy of a system is the infinitesimal transfer of heat to a closed system driving a reversible process, divided by the equilibrium temperature of the system. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/gibbs.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] / Free enthalpy |Useful work obtainable from a system at '''isobaric''' and '''isothermal''' conditions. Since H is U + pV, it can be replaced in the equation, making G = H - TS. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/deltaG.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] variation. |If ΔG < 0, the system's transformation can be spontaneous, if ΔG = 0 the transformation is inversible and the system is in an equilibrium state, if ΔG > 0 it can't be spontaneous. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/density_ideal.png |[https://en.wikipedia.org/wiki/Density#Changes_of_density Density] of an ideal gas. |M is molar mass. This means that the density of an ideal gas can be doubled by doubling the pressure, or by halving the absolute temperature. |} 14bdbb93800f864ef91c0af94a3f207e43a8e570 290 289 2012-05-01T19:43:48Z Vincent 1 /* List of equations */ adding dH(S,p) wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] or [https://en.wikipedia.org/wiki/Heat_capacity_ratio Heat capacity ratio] (adiabatic process in thermodynamics) | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | ρ (rho) | [https://en.wikipedia.org/wiki/Density Density] | kg.m^-3 (kg per cubic meter) |- | C, C_p, C_V | [https://en.wikipedia.org/wiki/Heat_capacity#Metrology Heat capacity], general, at constant pressure, at constant volume. | J.K^-1 (Joule per Kelvin) |- | G | [https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] | J (Joule) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: total energy of a thermodynamic system. | J (Joule) |- | ΔH_vap or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | M | [https://en.wikipedia.org/wiki/Mach_number Mach number] | no unit |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature. T₀ or T_t is the [https://en.wikipedia.org/wiki/Stagnation_temperature stagnation temperature]. | K (Kelvin) |- | S | [https://en.wikipedia.org/wiki/Entropy Entropy] | J.K^-1 (Joule per Kelvin) |- | U | [https://en.wikipedia.org/wiki/Internal_energy Internal energy] of a system (see first law of Thermodynamics below) | J (Joule) |- | V | Volume | m³ (cubic meter) |- | W | [https://en.wikipedia.org/wiki/Work_(thermodynamics) Work]: mechanical constraints on the system. | J (Joule) |- | a | [https://en.wikipedia.org/wiki/Speed_of_sound Speed of sound] in medium (used to calculate Mach number) | m.s^-1 |- | c | Velocity of a flow in thermodynamics, also noted V; generally noted u in fluid dynamics. | m.s^-1 |- | n | Quantity of matter | mol (mole) |- | p | Pressure. p_t is the [https://en.wikipedia.org/wiki/Stagnation_pressure stagnation pressure]. | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/QeqmL.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. Delta H equals Q only when pressure is constant (isobaric). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dUeqdQmindW.png |[https://en.wikipedia.org/wiki/First_law_of_thermodynamics First law of thermodynamics] |Variations of internal energy of a system between two states is the sum of the received heat and work (minus the ''given'' work). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/enthalpy.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] |Total amount of energy of a system, defined as the sum of the internal energy U of the system and pressure * volume at the boundary of the system and its environment. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/workExpand.png |Work of gas expansion. |Work done by expanding an ideal gas. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/entropy_dueqtdsmpdv.png |[https://en.wikipedia.org/wiki/Internal_energy Internal energy] change related to [https://en.wikipedia.org/wiki/Entropy entropy] |Internal energy related to entropy variation for a closed system in thermal equilibrium ([https://en.wikipedia.org/wiki/Fundamental_thermodynamic_relation fundamental thermodynamic relation]). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dheqtds.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] change |Enthalpy change depending on entropy and pressure changes, equation created from the mix of the basic ones above. |- |ΔS_universe = ΔS_surroundings + ΔS_system |Entropy variation as a whole. |Entropy variation of a system is generally compensated by the inverse variation of the surroundings, not including losses. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dS.png |[https://en.wikipedia.org/wiki/Second_law_of_thermodynamics Second law of thermodynamics] |A change in the entropy of a system is the infinitesimal transfer of heat to a closed system driving a reversible process, divided by the equilibrium temperature of the system. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/gibbs.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] / Free enthalpy |Useful work obtainable from a system at '''isobaric''' and '''isothermal''' conditions. Since H is U + pV, it can be replaced in the equation, making G = H - TS. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/deltaG.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] variation. |If ΔG < 0, the system's transformation can be spontaneous, if ΔG = 0 the transformation is inversible and the system is in an equilibrium state, if ΔG > 0 it can't be spontaneous. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/density_ideal.png |[https://en.wikipedia.org/wiki/Density#Changes_of_density Density] of an ideal gas. |M is molar mass. This means that the density of an ideal gas can be doubled by doubling the pressure, or by halving the absolute temperature. |} b63b689e6121364fa288624667acac083e0fd72e 296 290 2012-05-28T01:27:40Z Vincent 1 /* List of constants */ adding G wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] or [https://en.wikipedia.org/wiki/Heat_capacity_ratio Heat capacity ratio] (adiabatic process in thermodynamics) | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | ρ (rho) | [https://en.wikipedia.org/wiki/Density Density] | kg.m^-3 (kg per cubic meter) |- | C, C_p, C_V | [https://en.wikipedia.org/wiki/Heat_capacity#Metrology Heat capacity], general, at constant pressure, at constant volume. | J.K^-1 (Joule per Kelvin) |- | G | [https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] | J (Joule) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: total energy of a thermodynamic system. | J (Joule) |- | ΔH_vap or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | M | [https://en.wikipedia.org/wiki/Mach_number Mach number] | no unit |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature. T₀ or T_t is the [https://en.wikipedia.org/wiki/Stagnation_temperature stagnation temperature]. | K (Kelvin) |- | S | [https://en.wikipedia.org/wiki/Entropy Entropy] | J.K^-1 (Joule per Kelvin) |- | U | [https://en.wikipedia.org/wiki/Internal_energy Internal energy] of a system (see first law of Thermodynamics below) | J (Joule) |- | V | Volume | m³ (cubic meter) |- | W | [https://en.wikipedia.org/wiki/Work_(thermodynamics) Work]: mechanical constraints on the system. | J (Joule) |- | a | [https://en.wikipedia.org/wiki/Speed_of_sound Speed of sound] in medium (used to calculate Mach number) | m.s^-1 |- | c | Velocity of a flow in thermodynamics, also noted V; generally noted u in fluid dynamics. | m.s^-1 |- | n | Quantity of matter | mol (mole) |- | p | Pressure. p_t is the [https://en.wikipedia.org/wiki/Stagnation_pressure stagnation pressure]. | Pa (Pascal) |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | G | [https://en.wikipedia.org/wiki/Gravitational_constant Gravitational constant] | 6.674 | m³.kg^-1.s^-2 |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/QeqmL.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. Delta H equals Q only when pressure is constant (isobaric). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dUeqdQmindW.png |[https://en.wikipedia.org/wiki/First_law_of_thermodynamics First law of thermodynamics] |Variations of internal energy of a system between two states is the sum of the received heat and work (minus the ''given'' work). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/enthalpy.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] |Total amount of energy of a system, defined as the sum of the internal energy U of the system and pressure * volume at the boundary of the system and its environment. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/workExpand.png |Work of gas expansion. |Work done by expanding an ideal gas. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/entropy_dueqtdsmpdv.png |[https://en.wikipedia.org/wiki/Internal_energy Internal energy] change related to [https://en.wikipedia.org/wiki/Entropy entropy] |Internal energy related to entropy variation for a closed system in thermal equilibrium ([https://en.wikipedia.org/wiki/Fundamental_thermodynamic_relation fundamental thermodynamic relation]). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dheqtds.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] change |Enthalpy change depending on entropy and pressure changes, equation created from the mix of the basic ones above. |- |ΔS_universe = ΔS_surroundings + ΔS_system |Entropy variation as a whole. |Entropy variation of a system is generally compensated by the inverse variation of the surroundings, not including losses. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dS.png |[https://en.wikipedia.org/wiki/Second_law_of_thermodynamics Second law of thermodynamics] |A change in the entropy of a system is the infinitesimal transfer of heat to a closed system driving a reversible process, divided by the equilibrium temperature of the system. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/gibbs.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] / Free enthalpy |Useful work obtainable from a system at '''isobaric''' and '''isothermal''' conditions. Since H is U + pV, it can be replaced in the equation, making G = H - TS. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/deltaG.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] variation. |If ΔG < 0, the system's transformation can be spontaneous, if ΔG = 0 the transformation is inversible and the system is in an equilibrium state, if ΔG > 0 it can't be spontaneous. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/density_ideal.png |[https://en.wikipedia.org/wiki/Density#Changes_of_density Density] of an ideal gas. |M is molar mass. This means that the density of an ideal gas can be doubled by doubling the pressure, or by halving the absolute temperature. |} 1d9a9661a720225747a1e2cfad520f9ead92e299 0 9 288 85 2012-04-19T00:49:53Z Vincent 1 new sensors links (camera, pitot tubes and compass) and HTTPS links wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [https://www.armadeus.com/ Armadeus], based on an ARM (FreeScale) processor. Moreover, it [https://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [https://xenomai.org/ Xenomai] Linux-based RTOS. Armadeus board integration has a [[Armadeus|dedicated page]]. Other interesting embedded computer boards: the [https://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1] and the [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20]. They do not include a FPGA but are cheaper, the first has a great temperature range, and the second is more powerful and smaller. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [https://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. ===Telemetry=== It seems that the 900MHz version of the ZigBee communication standard is able to transmit at around 100kbps up to 10km. Taken from th ArduPilot page: Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [https://www.adafruit.com/products/126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [https://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [https://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: quite common nowadays, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the theoretically computed strength the mechanical structure has to sustain. To chose a sensor, sparkfun wrote an [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and might be replaced by a camera sensor for low rotation rates. * Magnetometer (3D compass): lots of sensors exist too, for example the [https://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag] and [https://www.sparkfun.com/products/8128 SCP1000], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * Thermometer: for systems health monitoring, like engines temperature. * GPS if USAF and sensors allow it in flight altitude. * Camera: 8-bit data port if possible, like the TCM8230MD sensor. Some ARM processors (i.MX) feature the Camera/CMOS Sensor Interface (CSI) and hardware-accelerated processing or compression from this port. Horizon sensor is provided by a camera. * Pitot tubes even exist in stores (like [https://store.diydrones.com/Kit_MPXV7002DP_p/kt-mpxv7002dp-01.htm DIYDrones])! * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [https://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. 0c1cfdf26f33c69be07bcffca720ad274f9afe93 0 57 295 2012-05-27T15:59:46Z Vincent 1 guidelines creation wikitext text/x-wiki =Guidelines= This page defines the writing rules and the methodology that has to be applied when working on the LCAS project. If you want to help with the project, the best way is probably to contribute to the opened paths of research in your domain or in an easier way, help with information gathering on which the research is based. A 3D rendering or drawing of the plane/rocket system would help visitors to understand the project at the first look to the main page, so if you feel you can make it, please go ahead. # This project is a scientific approach to the problem of low-cost lightweight orbital system, not a political approach. The goal is thus to really produce something at some point with no preconceived idea. # Information available on this website is used as the building block for the research being made, it thus has to be correct and as precise as possible, and should contain references when possible. Unverified information can be stored for future assessment but should be mentioned as unverified. # The language used is English from England. Some American English terms may be used if they are found like this in the bibliography, like airfoil instead of aerofoil. # HTTPS links are preferred over HTTP links. 4d07e5638717f616917e4742b68f5f9e2a4d7a2b 0 43 297 242 2012-05-28T15:18:11Z Vincent 1 adding subsections for fan and turbine wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine acting as the power plant behind it. Their flows mix at the exhaust, inside the engine for low bypass ratio engines and outside, in atmosphere, for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a hot gas flow, which would eventually corrode or induce fatigue on it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. * The hybrid turbine uses both axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its linear course, while extracting some of its energy for shaft rotation work. * The air breathed by the fan would be the mixed flow of the traditional fresh intake and the turbine discharge, providing higher temperature to the fan flow. The main advantage of this hot air is the increase in speed of sound and the ability to drive the fan at transonic regime on the same spool than the compressor operating at transonic regime too. Multi-spool engines have been created to allow such features in classical engines. The engine will inevitably be longer because room has to be made for the flow-mixing section. This will make the engine slightly heavier too. ==Hybrid turbine== ''work in progress'' The reasons why gas turbines employ axial-turbine stages should be listed here. For turbojets this is obviously due to the fact that the hot gas stream at engine discharge is what gives thrust, so the path of the hot gas in the turbine should be kept as linear as possible. ==Fan operation== ''work in progress'' Main question: is the fan much less efficient when it blows hot air than cold air? Is this efficiency overtaken by the efficiency increase in transonic regime versus subsonic? Thrust depends on the mass flow rate of expelled material. Higher temperature means that the density is lower, and thus that the mass flow rate should be lower too. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000-series aluminium alloys. ==Example implementation: full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' bypass ratio (BPR) because since flows are mixed before the fan intake, there is no clear separation between flows of the fan and of the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is thus the ratio between the sum of the fan duct mass flow rate and the core's mass flow rate over the core's mass flow rate (M_fan + M_core) / M_core. Let's take an example turbofan engine with a 13cm fan and a 9cm core (and compressor) diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both side ((13²-9²)/9²), and the BPR is in fact 2.09 (13²/9²). A great advantage of our aft-fan engine design is that both the first stage(s) of the compressor and the fan can operate at '''transonic''' speeds. A blade or a fan is said having a [http://en.wikipedia.org/wiki/Transonic transonic] operation when the flow passing through it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, while their root is generally subsonic too. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound, which depends on the temperature of the air flow. A transonic operation allows higher compression ratios to be achieved, with a lower efficiency than subsonic operation as a drawback, around 5% less because of the drag induced by shock waves. As a result, our design allows both the compressor entry stage and the fan to operate at transonic speeds, while they have the same rotation speed (one spool shaft), at high altitudes. This is possible because of the temperature difference in air passing through both. The compressor will breathe fresh air, which can go down as low as -57°C in altitude. In this case, the sonic speed is 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. If we fix a 200°C rise of this air flow, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. Having a transonic speed operation allows higher the compression ratios. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor, it's not able to have a high overall compression ratio (CR). With a transonic-operating first stage, we hope to have at least 2.0 CR for it, instead of the maximum 1.6 CR in subsonic operation, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Higher efficiency of the fan will allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. [[File:Engine core and fan.jpg|300px]] [[File:Engine core and fan side.jpg|350px]] ''The above schematics may not reflect actual proportions and have transparency inconsistencies.'' That said, the engine's core is the green internal tube. The fan sketched inside it, at the front of the engine, is the compressor's first stage rotor. The stator and other stage are not represented, and will obviously be placed on the conic part behind it. The red part is the annular [[Turbofan:Combustors|combustor]], not easy to see clearly in these pictures. '''The turbine is not represented''' either, but an axial rotor should be placed right after the combustor, and the hybrid turbine will be milled in the shaft in the curved part after a second axial turbine guide vane. We can see that the discharge of the turbine is mixed with the inlet air of the aft-mounted fan. The fan duct is thus longer than in traditional front-mounted engines, but the nozzle can be closer to the fan, allowing to work on both core and fan flow. [[Category:Turbofan|Alternative design]] a0660f76b0eacd5e5be6ff0fd6803a23a1b49fbe 298 297 2012-05-28T15:33:52Z Vincent 1 /* Hybrid turbine */ wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine acting as the power plant behind it. Their flows mix at the exhaust, inside the engine for low bypass ratio engines and outside, in atmosphere, for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a hot gas flow, which would eventually corrode or induce fatigue on it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. * The hybrid turbine uses both axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its linear course, while extracting some of its energy for shaft rotation work. * The air breathed by the fan would be the mixed flow of the traditional fresh intake and the turbine discharge, providing higher temperature to the fan flow. The main advantage of this hot air is the increase in speed of sound and the ability to drive the fan at transonic regime on the same spool than the compressor operating at transonic regime too. Multi-spool engines have been created to allow such features in classical engines. The engine will inevitably be longer because room has to be made for the flow-mixing section. This will make the engine slightly heavier too. ==Hybrid turbine== ''work in progress'' The reasons why gas turbines employ axial-turbine stages should be listed here. For turbojets this is obviously due to the fact that the hot gas stream at engine discharge is what gives thrust, so the path of the hot gas in the turbine should be kept as linear as possible. For turbofans, the core engine exhaust also plays a role in providing thrust, generally around 10% of the total engine's thrust. Diverting the hot gas flow results in heavy loss of kinetic energy. Passing through a centrifugal-turbine will likely extract lots of energy from the gas and not allow further work to be done from it. It also means that a centrifugal-turbine has to cope with much higher temperatures than axial-turbines, which can be a real show-stopper since axial-turbines already have important issues with temperature. In our case, we need to divert the hot gas flow in order to have it mixed properly with the cold air from the bypass. The diversion thus doesn't have to be perpendicular to the flow and the turbine will be operating in a hybrid-mode, with a centrifugal-mode upstream and axial-mode downstream, after the flow has been diverted. This will result in a turbine having boomerang-shaped blades, as depicted below. ==Fan operation== ''work in progress'' Main question: is the fan much less efficient when it blows hot air than cold air? Is this efficiency overtaken by the efficiency increase in transonic regime versus subsonic? Thrust depends on the mass flow rate of expelled material. Higher temperature means that the density is lower, and thus that the mass flow rate should be lower too. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000-series aluminium alloys. ==Example implementation: full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' bypass ratio (BPR) because since flows are mixed before the fan intake, there is no clear separation between flows of the fan and of the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is thus the ratio between the sum of the fan duct mass flow rate and the core's mass flow rate over the core's mass flow rate (M_fan + M_core) / M_core. Let's take an example turbofan engine with a 13cm fan and a 9cm core (and compressor) diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both side ((13²-9²)/9²), and the BPR is in fact 2.09 (13²/9²). A great advantage of our aft-fan engine design is that both the first stage(s) of the compressor and the fan can operate at '''transonic''' speeds. A blade or a fan is said having a [http://en.wikipedia.org/wiki/Transonic transonic] operation when the flow passing through it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, while their root is generally subsonic too. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound, which depends on the temperature of the air flow. A transonic operation allows higher compression ratios to be achieved, with a lower efficiency than subsonic operation as a drawback, around 5% less because of the drag induced by shock waves. As a result, our design allows both the compressor entry stage and the fan to operate at transonic speeds, while they have the same rotation speed (one spool shaft), at high altitudes. This is possible because of the temperature difference in air passing through both. The compressor will breathe fresh air, which can go down as low as -57°C in altitude. In this case, the sonic speed is 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. If we fix a 200°C rise of this air flow, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. Having a transonic speed operation allows higher the compression ratios. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor, it's not able to have a high overall compression ratio (CR). With a transonic-operating first stage, we hope to have at least 2.0 CR for it, instead of the maximum 1.6 CR in subsonic operation, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Higher efficiency of the fan will allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. [[File:Engine core and fan.jpg|300px]] [[File:Engine core and fan side.jpg|350px]] ''The above schematics may not reflect actual proportions and have transparency inconsistencies.'' That said, the engine's core is the green internal tube. The fan sketched inside it, at the front of the engine, is the compressor's first stage rotor. The stator and other stage are not represented, and will obviously be placed on the conic part behind it. The red part is the annular [[Turbofan:Combustors|combustor]], not easy to see clearly in these pictures. '''The turbine is not represented''' either, but an axial rotor should be placed right after the combustor, and the hybrid turbine will be milled in the shaft in the curved part after a second axial turbine guide vane. We can see that the discharge of the turbine is mixed with the inlet air of the aft-mounted fan. The fan duct is thus longer than in traditional front-mounted engines, but the nozzle can be closer to the fan, allowing to work on both core and fan flow. [[Category:Turbofan|Alternative design]] 817cd46dbd66aacf463259597a4eeffcf67d3f30 301 298 2012-05-28T18:58:32Z Vincent 1 /* Hybrid turbine */ wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine acting as the power plant behind it. Their flows mix at the exhaust, inside the engine for low bypass ratio engines and outside, in atmosphere, for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a hot gas flow, which would eventually corrode or induce fatigue on it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. * The hybrid turbine uses both axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its linear course, while extracting some of its energy for shaft rotation work. * The air breathed by the fan would be the mixed flow of the traditional fresh intake and the turbine discharge, providing higher temperature to the fan flow. The main advantage of this hot air is the increase in speed of sound and the ability to drive the fan at transonic regime on the same spool than the compressor operating at transonic regime too. Multi-spool engines have been created to allow such features in classical engines. The engine will inevitably be longer because room has to be made for the flow-mixing section. This will make the engine slightly heavier too. ==Hybrid turbine== ''work in progress'' The reasons why gas turbines employ axial-turbine stages should be listed here. For turbojets this is obviously due to the fact that the hot gas stream at engine discharge is what gives thrust, so the path of the hot gas in the turbine should be kept as linear as possible. For turbofans, the core engine exhaust also plays a role in providing thrust, generally around 10% of the total engine's thrust. Diverting the hot gas flow results in heavy loss of kinetic energy. Passing through a centrifugal-turbine will likely extract lots of energy from the gas and not allow further work to be done from it. It also means that a centrifugal-turbine has to cope with much higher temperatures than axial-turbines, which can be a real show-stopper since axial-turbines already have important issues with temperature. In our case, we need to divert the hot gas flow in order to have it mixed properly with the cold air from the bypass. The diversion thus doesn't have to be more than 45° and the turbine will be operating in a hybrid-mode, with a centrifugal-mode upstream and axial-mode downstream, after the flow has been diverted. This will result in a turbine having boomerang-shaped blades, as depicted below. [[File:Hybrid_turbine.png]] ==Fan operation== ''work in progress'' Main question: is the fan much less efficient when it blows hot air than cold air? Is this efficiency overtaken by the efficiency increase in transonic regime versus subsonic? Thrust depends on the mass flow rate of expelled material. Higher temperature means that the density is lower, and thus that the mass flow rate should be lower too. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000-series aluminium alloys. ==Example implementation: full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' bypass ratio (BPR) because since flows are mixed before the fan intake, there is no clear separation between flows of the fan and of the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is thus the ratio between the sum of the fan duct mass flow rate and the core's mass flow rate over the core's mass flow rate (M_fan + M_core) / M_core. Let's take an example turbofan engine with a 13cm fan and a 9cm core (and compressor) diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both side ((13²-9²)/9²), and the BPR is in fact 2.09 (13²/9²). A great advantage of our aft-fan engine design is that both the first stage(s) of the compressor and the fan can operate at '''transonic''' speeds. A blade or a fan is said having a [http://en.wikipedia.org/wiki/Transonic transonic] operation when the flow passing through it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, while their root is generally subsonic too. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound, which depends on the temperature of the air flow. A transonic operation allows higher compression ratios to be achieved, with a lower efficiency than subsonic operation as a drawback, around 5% less because of the drag induced by shock waves. As a result, our design allows both the compressor entry stage and the fan to operate at transonic speeds, while they have the same rotation speed (one spool shaft), at high altitudes. This is possible because of the temperature difference in air passing through both. The compressor will breathe fresh air, which can go down as low as -57°C in altitude. In this case, the sonic speed is 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. If we fix a 200°C rise of this air flow, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. Having a transonic speed operation allows higher the compression ratios. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor, it's not able to have a high overall compression ratio (CR). With a transonic-operating first stage, we hope to have at least 2.0 CR for it, instead of the maximum 1.6 CR in subsonic operation, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Higher efficiency of the fan will allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. [[File:Engine core and fan.jpg|300px]] [[File:Engine core and fan side.jpg|350px]] ''The above schematics may not reflect actual proportions and have transparency inconsistencies.'' That said, the engine's core is the green internal tube. The fan sketched inside it, at the front of the engine, is the compressor's first stage rotor. The stator and other stage are not represented, and will obviously be placed on the conic part behind it. The red part is the annular [[Turbofan:Combustors|combustor]], not easy to see clearly in these pictures. '''The turbine is not represented''' either, but an axial rotor should be placed right after the combustor, and the hybrid turbine will be milled in the shaft in the curved part after a second axial turbine guide vane. We can see that the discharge of the turbine is mixed with the inlet air of the aft-mounted fan. The fan duct is thus longer than in traditional front-mounted engines, but the nozzle can be closer to the fan, allowing to work on both core and fan flow. [[Category:Turbofan|Alternative design]] 7b4857f8557d7fe7e21aec6289a4afcf35fe941e 6 59 300 2012-05-28T18:54:41Z Vincent 1 Hybrid turbine sketch wikitext text/x-wiki Hybrid turbine sketch 7ade73c50e2c15a14be1c868607b4580a2fc4084 6 59 302 300 2012-05-28T19:03:37Z Vincent 1 uploaded a new version of "[[File:Hybrid turbine.png]]":&#32;background wikitext text/x-wiki Hybrid turbine sketch 7ade73c50e2c15a14be1c868607b4580a2fc4084 303 302 2012-05-28T19:18:03Z Vincent 1 uploaded a new version of "[[File:Hybrid turbine.png]]":&#32;changing colors for background and flows wikitext text/x-wiki Hybrid turbine sketch 7ade73c50e2c15a14be1c868607b4580a2fc4084 306 303 2012-05-28T22:58:14Z Vincent 1 wikitext text/x-wiki Hybrid turbine sketch. This depicts the ending part of our turbofan engine in a upper longitudinal cut, from the combustor discharge to the nozzle entry. The stator is up, the rotor is down, cut at its axis. 2b1fb45a12ac74cd09d7e97d3cd9340af2012e95 312 306 2012-05-31T23:38:39Z Vincent 1 uploaded a new version of "[[File:Hybrid turbine.png]]":&#32;glimpses of an expander and minor improvements wikitext text/x-wiki Hybrid turbine sketch. This depicts the ending part of our turbofan engine in a upper longitudinal cut, from the combustor discharge to the nozzle entry. The stator is up, the rotor is down, cut at its axis. 2b1fb45a12ac74cd09d7e97d3cd9340af2012e95 6 60 304 2012-05-28T22:28:01Z Vincent 1 T-S diagram for the mixed/hybrid turbofan. wikitext text/x-wiki T-S diagram for the mixed/hybrid turbofan. 577abc40c397b46de550d49516764389112de5d0 0 43 305 301 2012-05-28T22:30:43Z Vincent 1 thermodynamics wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine acting as the power plant behind it. Their flows mix at the exhaust, inside the engine for low bypass ratio engines and outside, in atmosphere, for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a hot gas flow, which would eventually corrode or induce fatigue on it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. * The hybrid turbine uses both axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its linear course, while extracting some of its energy for shaft rotation work. * The air breathed by the fan would be the mixed flow of the traditional fresh intake and the turbine discharge, providing higher temperature to the fan flow. The main advantage of this hot air is the increase in speed of sound and the ability to drive the fan at transonic regime on the same spool than the compressor operating at transonic regime too. Multi-spool engines have been created to allow such features in classical engines. The engine will inevitably be longer because room has to be made for the flow-mixing section. This will make the engine slightly heavier too. ==Hybrid turbine== ''work in progress'' The reasons why gas turbines employ axial-turbine stages should be listed here. For turbojets this is obviously due to the fact that the hot gas stream at engine discharge is what gives thrust, so the path of the hot gas in the turbine should be kept as linear as possible. For turbofans, the core engine exhaust also plays a role in providing thrust, generally around 10% of the total engine's thrust. Diverting the hot gas flow results in heavy loss of kinetic energy. Passing through a centrifugal-turbine will likely extract lots of energy from the gas and not allow further work to be done from it. It also means that a centrifugal-turbine has to cope with much higher temperatures than axial-turbines, which can be a real show-stopper since axial-turbines already have important issues with temperature. In our case, we need to divert the hot gas flow in order to have it mixed properly with the cold air from the bypass. The diversion thus doesn't have to be more than 45° and the turbine will be operating in a hybrid-mode, with a centrifugal-mode upstream and axial-mode downstream, after the flow has been diverted. This will result in a turbine having boomerang-shaped blades, as depicted below. [[File:Hybrid_turbine.png]] ==Fan operation== ''work in progress'' Main question: is the fan much less efficient when it blows hot air than cold air? Is this efficiency overtaken by the efficiency increase in transonic regime versus subsonic? Thrust depends on the mass flow rate of expelled material. Higher temperature means that the density is lower, and thus that the mass flow rate should be lower too. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000-series aluminium alloys. ==Thermodynamic analysis== Turbine engines are assimilated to a [https://en.wikipedia.org/wiki/Brayton_cycle Brayton cycle]. It consists of 4 processes: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion in two steps - first made by the turbine, the extracted work is used to drive both the compressor and the payload. In the case of a turbofan the payload is the fan. Then, the nozzle finishes the expansion and exhausts gas at atmosphere pressure. # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. In our approach, the expansion has to be considered as two different steps because there is a heat exchange between the two. Also, the fan doesn't operate in an adiabatic process since there is a heat input made by the core engine exhaust gas mixing. The two cycles, the core engine cycle and the fan cycle are now combined. The new process are: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion - made by the turbine ## isobaric heat rejection - made by the flow mixing ## adiabatic compression - made by the fan # adiabatic expansion - made by the nozzle # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. [[File:Thermodynamic_diagrams.png|center|250px]] ==Example implementation: full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' bypass ratio (BPR) because since flows are mixed before the fan intake, there is no clear separation between flows of the fan and of the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is thus the ratio between the sum of the fan duct mass flow rate and the core's mass flow rate over the core's mass flow rate (M_fan + M_core) / M_core. Let's take an example turbofan engine with a 13cm fan and a 9cm core (and compressor) diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both side ((13²-9²)/9²), and the BPR is in fact 2.09 (13²/9²). A great advantage of our aft-fan engine design is that both the first stage(s) of the compressor and the fan can operate at '''transonic''' speeds. A blade or a fan is said having a [http://en.wikipedia.org/wiki/Transonic transonic] operation when the flow passing through it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, while their root is generally subsonic too. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound, which depends on the temperature of the air flow. A transonic operation allows higher compression ratios to be achieved, with a lower efficiency than subsonic operation as a drawback, around 5% less because of the drag induced by shock waves. As a result, our design allows both the compressor entry stage and the fan to operate at transonic speeds, while they have the same rotation speed (one spool shaft), at high altitudes. This is possible because of the temperature difference in air passing through both. The compressor will breathe fresh air, which can go down as low as -57°C in altitude. In this case, the sonic speed is 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. If we fix a 200°C rise of this air flow, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. Having a transonic speed operation allows higher the compression ratios. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor, it's not able to have a high overall compression ratio (CR). With a transonic-operating first stage, we hope to have at least 2.0 CR for it, instead of the maximum 1.6 CR in subsonic operation, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Higher efficiency of the fan will allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. [[File:Engine core and fan.jpg|300px]] [[File:Engine core and fan side.jpg|350px]] ''The above schematics may not reflect actual proportions and have transparency inconsistencies.'' That said, the engine's core is the green internal tube. The fan sketched inside it, at the front of the engine, is the compressor's first stage rotor. The stator and other stage are not represented, and will obviously be placed on the conic part behind it. The red part is the annular [[Turbofan:Combustors|combustor]], not easy to see clearly in these pictures. '''The turbine is not represented''' either, but an axial rotor should be placed right after the combustor, and the hybrid turbine will be milled in the shaft in the curved part after a second axial turbine guide vane. We can see that the discharge of the turbine is mixed with the inlet air of the aft-mounted fan. The fan duct is thus longer than in traditional front-mounted engines, but the nozzle can be closer to the fan, allowing to work on both core and fan flow. [[Category:Turbofan|Alternative design]] ebb22c395a1238fc8bbb7f159fb0b427d45658d1 310 305 2012-05-31T18:08:18Z Vincent 1 /* Hybrid turbine */ wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine acting as the power plant behind it. Their flows mix at the exhaust, inside the engine for low bypass ratio engines and outside, in atmosphere, for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a hot gas flow, which would eventually corrode or induce fatigue on it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, long ago. From these two facts, we propose a novel design for turbofans: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. * The hybrid turbine uses both axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its linear course, while extracting some of its energy for shaft rotation work. * The air breathed by the fan would be the mixed flow of the traditional fresh intake and the turbine discharge, providing higher temperature to the fan flow. The main advantage of this hot air is the increase in speed of sound and the ability to drive the fan at transonic regime on the same spool than the compressor operating at transonic regime too. Multi-spool engines have been created to allow such features in classical engines. The engine will inevitably be longer because room has to be made for the flow-mixing section. This will make the engine slightly heavier too. ==Hybrid turbine== ''work in progress'' ''The reasons why gas turbines employ axial-turbine stages should be listed here''. For turbojets this is obviously due to the fact that the hot gas stream at engine discharge is what gives thrust, so the path of the hot gas from the turbine to the nozzle should be kept as linear as possible. For turbofans, the core engine exhaust also plays a role in providing thrust, generally around 10% of the total engine's thrust. Diverting the hot gas flow results in heavy loss of kinetic energy. Passing through a centrifugal-turbine extracts lots of energy from the gas and does not allow further work to be done from it. It also means that a centrifugal-turbine has to cope with much higher temperatures than axial-turbines, which can be a real show-stopper since axial-turbines already have important issues with temperature. In our case, we need to divert the hot gas flow in order to have it mixed properly with the cold air from the bypass. The diversion thus should be less than 45° and the turbine will be operating in a hybrid-mode, with a centrifugal-mode upstream and axial-mode downstream, harvesting energy from the diverted flow. This will result in a turbine having boomerang-shaped blades, as the green part depicted below. [[File:Hybrid_turbine.png]] ==Fan operation== ''work in progress'' Main question: is the fan much less efficient when it blows hot air than cold air? Is this efficiency overtaken by the efficiency increase in transonic regime versus subsonic? Thrust depends on the mass flow rate of expelled material. Higher temperature means that the density is lower, and thus that the mass flow rate should be lower too. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000-series aluminium alloys. ==Thermodynamic analysis== Turbine engines are assimilated to a [https://en.wikipedia.org/wiki/Brayton_cycle Brayton cycle]. It consists of 4 processes: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion in two steps - first made by the turbine, the extracted work is used to drive both the compressor and the payload. In the case of a turbofan the payload is the fan. Then, the nozzle finishes the expansion and exhausts gas at atmosphere pressure. # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. In our approach, the expansion has to be considered as two different steps because there is a heat exchange between the two. Also, the fan doesn't operate in an adiabatic process since there is a heat input made by the core engine exhaust gas mixing. The two cycles, the core engine cycle and the fan cycle are now combined. The new process are: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion - made by the turbine ## isobaric heat rejection - made by the flow mixing ## adiabatic compression - made by the fan # adiabatic expansion - made by the nozzle # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. [[File:Thermodynamic_diagrams.png|center|250px]] ==Example implementation: full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' bypass ratio (BPR) because since flows are mixed before the fan intake, there is no clear separation between flows of the fan and of the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is thus the ratio between the sum of the fan duct mass flow rate and the core's mass flow rate over the core's mass flow rate (M_fan + M_core) / M_core. Let's take an example turbofan engine with a 13cm fan and a 9cm core (and compressor) diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both side ((13²-9²)/9²), and the BPR is in fact 2.09 (13²/9²). A great advantage of our aft-fan engine design is that both the first stage(s) of the compressor and the fan can operate at '''transonic''' speeds. A blade or a fan is said having a [http://en.wikipedia.org/wiki/Transonic transonic] operation when the flow passing through it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, while their root is generally subsonic too. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound, which depends on the temperature of the air flow. A transonic operation allows higher compression ratios to be achieved, with a lower efficiency than subsonic operation as a drawback, around 5% less because of the drag induced by shock waves. As a result, our design allows both the compressor entry stage and the fan to operate at transonic speeds, while they have the same rotation speed (one spool shaft), at high altitudes. This is possible because of the temperature difference in air passing through both. The compressor will breathe fresh air, which can go down as low as -57°C in altitude. In this case, the sonic speed is 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. If we fix a 200°C rise of this air flow, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. Having a transonic speed operation allows higher the compression ratios. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor, it's not able to have a high overall compression ratio (CR). With a transonic-operating first stage, we hope to have at least 2.0 CR for it, instead of the maximum 1.6 CR in subsonic operation, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Higher efficiency of the fan will allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. [[File:Engine core and fan.jpg|300px]] [[File:Engine core and fan side.jpg|350px]] ''The above schematics may not reflect actual proportions and have transparency inconsistencies.'' That said, the engine's core is the green internal tube. The fan sketched inside it, at the front of the engine, is the compressor's first stage rotor. The stator and other stage are not represented, and will obviously be placed on the conic part behind it. The red part is the annular [[Turbofan:Combustors|combustor]], not easy to see clearly in these pictures. '''The turbine is not represented''' either, but an axial rotor should be placed right after the combustor, and the hybrid turbine will be milled in the shaft in the curved part after a second axial turbine guide vane. We can see that the discharge of the turbine is mixed with the inlet air of the aft-mounted fan. The fan duct is thus longer than in traditional front-mounted engines, but the nozzle can be closer to the fan, allowing to work on both core and fan flow. [[Category:Turbofan|Alternative design]] 0ce00e7826a7268375b9007d5d2c713c02d39ea0 311 310 2012-05-31T18:42:38Z Vincent 1 /* Alternative design for turbofans */ wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine acting as the power plant behind it. Their flows mix at the exhaust, inside the engine for low bypass ratio engines and outside, in atmosphere, for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a hot gas flow, which would eventually corrode or induce fatigue on it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, [https://en.wikipedia.org/wiki/Heinkel_HeS_1 the HeS 1] ([http://airandspace.si.edu/collections/artifact.cfm?id=A19810039000 picture]), the first that actually flew a plane. From these two facts, we propose a novel design for turbofans: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. * The hybrid turbine uses both axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its linear course, while extracting some of its energy for shaft rotation work. The centrifugal part is crossed by the flow from the centre to the outer, contrary to what has been done in the HeS 1. * The air breathed by the fan would be the mixed flow of the traditional fresh intake and the turbine discharge, providing higher temperature to the fan flow. The main advantage of this hot air is the increase in speed of sound and the ability to drive the fan at transonic regime on the same spool than the compressor operating at transonic regime too. Multi-spool engines have been created to allow such features in classical engines. The engine will inevitably be longer because room has to be made for the flow-mixing section. This will make the engine slightly heavier too. ==Hybrid turbine== ''work in progress'' ''The reasons why gas turbines employ axial-turbine stages should be listed here''. For turbojets this is obviously due to the fact that the hot gas stream at engine discharge is what gives thrust, so the path of the hot gas from the turbine to the nozzle should be kept as linear as possible. For turbofans, the core engine exhaust also plays a role in providing thrust, generally around 10% of the total engine's thrust. Diverting the hot gas flow results in heavy loss of kinetic energy. Passing through a centrifugal-turbine extracts lots of energy from the gas and does not allow further work to be done from it. It also means that a centrifugal-turbine has to cope with much higher temperatures than axial-turbines, which can be a real show-stopper since axial-turbines already have important issues with temperature. In our case, we need to divert the hot gas flow in order to have it mixed properly with the cold air from the bypass. The diversion thus should be less than 45° and the turbine will be operating in a hybrid-mode, with a centrifugal-mode upstream and axial-mode downstream, harvesting energy from the diverted flow. This will result in a turbine having boomerang-shaped blades, as the green part depicted below. [[File:Hybrid_turbine.png]] ==Bypass flow and core discharge flow mixing== The hybrid turbine will be the last part crossed by the flow before the mixing. The hot flow will be abruptly expanded in the mixing area, increasing the static pressure and the temperature. This will allow for an even larger heating of the bypass flow. The static pressure increase will ''hopefully'' be the counterpart to air density decrease due to the temperature increase, still allowing the fan to operate in an efficient way. ''This has to be verified.'' An exhaust diffuser should be designed to control the expansion and the mixing. ==Fan operation== ''work in progress'' Main question: is the fan much less efficient when it blows hot air than cold air? Is this efficiency overtaken by the efficiency increase in transonic regime versus subsonic? Thrust depends on the mass flow rate of expelled material. Higher temperature means that the density is lower, and thus that the mass flow rate should be lower too. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000-series aluminium alloys. ==Thermodynamic analysis== Turbine engines are assimilated to a [https://en.wikipedia.org/wiki/Brayton_cycle Brayton cycle]. It consists of 4 processes: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion in two steps - first made by the turbine, the extracted work is used to drive both the compressor and the payload. In the case of a turbofan the payload is the fan. Then, the nozzle finishes the expansion and exhausts gas at atmosphere pressure. # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. In our approach, the expansion has to be considered as two different steps because there is a heat exchange between the two. Also, the fan doesn't operate in an adiabatic process since there is a heat input made by the core engine exhaust gas mixing. The two cycles, the core engine cycle and the fan cycle are now combined. The new process are: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion - made by the turbine ## isobaric heat rejection - made by the flow mixing ## adiabatic compression - made by the fan # adiabatic expansion - made by the nozzle # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. [[File:Thermodynamic_diagrams.png|center|250px]] ==Example implementation: full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' bypass ratio (BPR) because since flows are mixed before the fan intake, there is no clear separation between flows of the fan and of the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is thus the ratio between the sum of the fan duct mass flow rate and the core's mass flow rate over the core's mass flow rate (M_fan + M_core) / M_core. Let's take an example turbofan engine with a 13cm fan and a 9cm core (and compressor) diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both side ((13²-9²)/9²), and the BPR is in fact 2.09 (13²/9²). A great advantage of our aft-fan engine design is that both the first stage(s) of the compressor and the fan can operate at '''transonic''' speeds. A blade or a fan is said having a [http://en.wikipedia.org/wiki/Transonic transonic] operation when the flow passing through it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, while their root is generally subsonic too. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound, which depends on the temperature of the air flow. A transonic operation allows higher compression ratios to be achieved, with a lower efficiency than subsonic operation as a drawback, around 5% less because of the drag induced by shock waves. As a result, our design allows both the compressor entry stage and the fan to operate at transonic speeds, while they have the same rotation speed (one spool shaft), at high altitudes. This is possible because of the temperature difference in air passing through both. The compressor will breathe fresh air, which can go down as low as -57°C in altitude. In this case, the sonic speed is 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. If we fix a 200°C rise of this air flow, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. Having a transonic speed operation allows higher the compression ratios. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor, it's not able to have a high overall compression ratio (CR). With a transonic-operating first stage, we hope to have at least 2.0 CR for it, instead of the maximum 1.6 CR in subsonic operation, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Higher efficiency of the fan will allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. [[File:Engine core and fan.jpg|300px]] [[File:Engine core and fan side.jpg|350px]] ''The above schematics may not reflect actual proportions and have transparency inconsistencies.'' That said, the engine's core is the green internal tube. The fan sketched inside it, at the front of the engine, is the compressor's first stage rotor. The stator and other stage are not represented, and will obviously be placed on the conic part behind it. The red part is the annular [[Turbofan:Combustors|combustor]], not easy to see clearly in these pictures. '''The turbine is not represented''' either, but an axial rotor should be placed right after the combustor, and the hybrid turbine will be milled in the shaft in the curved part after a second axial turbine guide vane. We can see that the discharge of the turbine is mixed with the inlet air of the aft-mounted fan. The fan duct is thus longer than in traditional front-mounted engines, but the nozzle can be closer to the fan, allowing to work on both core and fan flow. [[Category:Turbofan|Alternative design]] 957cbafeb77ce8c712935c4eda29e8b4aa4f6e15 313 311 2012-06-01T01:56:23Z Vincent 1 /* Bypass flow and core discharge flow mixing */ wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine acting as the power plant behind it. Their flows mix at the exhaust, inside the engine for low bypass ratio engines and outside, in atmosphere, for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a hot gas flow, which would eventually corrode or induce fatigue on it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, [https://en.wikipedia.org/wiki/Heinkel_HeS_1 the HeS 1] ([http://airandspace.si.edu/collections/artifact.cfm?id=A19810039000 picture]), the first that actually flew a plane. From these two facts, we propose a novel design for turbofans: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. * The hybrid turbine uses both axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its linear course, while extracting some of its energy for shaft rotation work. The centrifugal part is crossed by the flow from the centre to the outer, contrary to what has been done in the HeS 1. * The air breathed by the fan would be the mixed flow of the traditional fresh intake and the turbine discharge, providing higher temperature to the fan flow. The main advantage of this hot air is the increase in speed of sound and the ability to drive the fan at transonic regime on the same spool than the compressor operating at transonic regime too. Multi-spool engines have been created to allow such features in classical engines. The engine will inevitably be longer because room has to be made for the flow-mixing section. This will make the engine slightly heavier too. ==Hybrid turbine== ''work in progress'' ''The reasons why gas turbines employ axial-turbine stages should be listed here''. For turbojets this is obviously due to the fact that the hot gas stream at engine discharge is what gives thrust, so the path of the hot gas from the turbine to the nozzle should be kept as linear as possible. For turbofans, the core engine exhaust also plays a role in providing thrust, generally around 10% of the total engine's thrust. Diverting the hot gas flow results in heavy loss of kinetic energy. Passing through a centrifugal-turbine extracts lots of energy from the gas and does not allow further work to be done from it. It also means that a centrifugal-turbine has to cope with much higher temperatures than axial-turbines, which can be a real show-stopper since axial-turbines already have important issues with temperature. In our case, we need to divert the hot gas flow in order to have it mixed properly with the cold air from the bypass. The diversion thus should be less than 45° and the turbine will be operating in a hybrid-mode, with a centrifugal-mode upstream and axial-mode downstream, harvesting energy from the diverted flow. This will result in a turbine having boomerang-shaped blades, as the green part depicted below. [[File:Hybrid_turbine.png]] ==Bypass flow and core discharge flow mixing== ''work in progress'' The hybrid turbine will be the last part crossed by the flow before the mixing. The hot flow will be abruptly expanded in the mixing area, increasing the static pressure (and the temperature? if it's an isentropic expansion it's a decrease of temperature, if it's a conservation of momentum as in continuum mechanics it's an increase I believe. This will allow for an even larger heating of the bypass flow.) The static pressure increase will ''hopefully'' be the counterpart to air density decrease due to the temperature increase, still allowing the fan to operate in an efficient way. ''This has to be verified.'' An exhaust diffuser should be designed to control the expansion and the mixing. The mixing can be seen as a '''preheat''', like the reheat but before the fan has done its compression work, allowing the nozzle to accelerate further the exhaust gas and thus providing a higher mass flow rate. ==Fan operation== ''work in progress'' Main question: is the fan much less efficient when it blows hot air than cold air? Is this efficiency overtaken by the efficiency increase in transonic regime versus subsonic? Thrust depends on the mass flow rate of expelled material. Higher temperature means that the density is lower, and thus that the mass flow rate should be lower too. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000-series aluminium alloys. ==Thermodynamic analysis== Turbine engines are assimilated to a [https://en.wikipedia.org/wiki/Brayton_cycle Brayton cycle]. It consists of 4 processes: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion in two steps - first made by the turbine, the extracted work is used to drive both the compressor and the payload. In the case of a turbofan the payload is the fan. Then, the nozzle finishes the expansion and exhausts gas at atmosphere pressure. # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. In our approach, the expansion has to be considered as two different steps because there is a heat exchange between the two. Also, the fan doesn't operate in an adiabatic process since there is a heat input made by the core engine exhaust gas mixing. The two cycles, the core engine cycle and the fan cycle are now combined. The new process are: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion - made by the turbine ## isobaric heat rejection - made by the flow mixing ## adiabatic compression - made by the fan # adiabatic expansion - made by the nozzle # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. [[File:Thermodynamic_diagrams.png|center|250px]] ==Example implementation: full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' bypass ratio (BPR) because since flows are mixed before the fan intake, there is no clear separation between flows of the fan and of the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is thus the ratio between the sum of the fan duct mass flow rate and the core's mass flow rate over the core's mass flow rate (M_fan + M_core) / M_core. Let's take an example turbofan engine with a 13cm fan and a 9cm core (and compressor) diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both side ((13²-9²)/9²), and the BPR is in fact 2.09 (13²/9²). A great advantage of our aft-fan engine design is that both the first stage(s) of the compressor and the fan can operate at '''transonic''' speeds. A blade or a fan is said having a [http://en.wikipedia.org/wiki/Transonic transonic] operation when the flow passing through it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, while their root is generally subsonic too. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound, which depends on the temperature of the air flow. A transonic operation allows higher compression ratios to be achieved, with a lower efficiency than subsonic operation as a drawback, around 5% less because of the drag induced by shock waves. As a result, our design allows both the compressor entry stage and the fan to operate at transonic speeds, while they have the same rotation speed (one spool shaft), at high altitudes. This is possible because of the temperature difference in air passing through both. The compressor will breathe fresh air, which can go down as low as -57°C in altitude. In this case, the sonic speed is 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. If we fix a 200°C rise of this air flow, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. Having a transonic speed operation allows higher the compression ratios. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor, it's not able to have a high overall compression ratio (CR). With a transonic-operating first stage, we hope to have at least 2.0 CR for it, instead of the maximum 1.6 CR in subsonic operation, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Higher efficiency of the fan will allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. [[File:Engine core and fan.jpg|300px]] [[File:Engine core and fan side.jpg|350px]] ''The above schematics may not reflect actual proportions and have transparency inconsistencies.'' That said, the engine's core is the green internal tube. The fan sketched inside it, at the front of the engine, is the compressor's first stage rotor. The stator and other stage are not represented, and will obviously be placed on the conic part behind it. The red part is the annular [[Turbofan:Combustors|combustor]], not easy to see clearly in these pictures. '''The turbine is not represented''' either, but an axial rotor should be placed right after the combustor, and the hybrid turbine will be milled in the shaft in the curved part after a second axial turbine guide vane. We can see that the discharge of the turbine is mixed with the inlet air of the aft-mounted fan. The fan duct is thus longer than in traditional front-mounted engines, but the nozzle can be closer to the fan, allowing to work on both core and fan flow. [[Category:Turbofan|Alternative design]] c042dab407f20386541fdea739009d33af347a3b 314 313 2012-06-02T02:12:04Z Vincent 1 /* Bypass flow and core discharge flow mixing */ wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine acting as the power plant behind it. Their flows mix at the exhaust, inside the engine for low bypass ratio engines and outside, in atmosphere, for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a hot gas flow, which would eventually corrode or induce fatigue on it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, [https://en.wikipedia.org/wiki/Heinkel_HeS_1 the HeS 1] ([http://airandspace.si.edu/collections/artifact.cfm?id=A19810039000 picture]), the first that actually flew a plane. From these two facts, we propose a novel design for turbofans: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. * The hybrid turbine uses both axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its linear course, while extracting some of its energy for shaft rotation work. The centrifugal part is crossed by the flow from the centre to the outer, contrary to what has been done in the HeS 1. * The air breathed by the fan would be the mixed flow of the traditional fresh intake and the turbine discharge, providing higher temperature to the fan flow. The main advantage of this hot air is the increase in speed of sound and the ability to drive the fan at transonic regime on the same spool than the compressor operating at transonic regime too. Multi-spool engines have been created to allow such features in classical engines. The engine will inevitably be longer because room has to be made for the flow-mixing section. This will make the engine slightly heavier too. ==Hybrid turbine== ''work in progress'' ''The reasons why gas turbines employ axial-turbine stages should be listed here''. For turbojets this is obviously due to the fact that the hot gas stream at engine discharge is what gives thrust, so the path of the hot gas from the turbine to the nozzle should be kept as linear as possible. For turbofans, the core engine exhaust also plays a role in providing thrust, generally around 10% of the total engine's thrust. Diverting the hot gas flow results in heavy loss of kinetic energy. Passing through a centrifugal-turbine extracts lots of energy from the gas and does not allow further work to be done from it. It also means that a centrifugal-turbine has to cope with much higher temperatures than axial-turbines, which can be a real show-stopper since axial-turbines already have important issues with temperature. In our case, we need to divert the hot gas flow in order to have it mixed properly with the cold air from the bypass. The diversion thus should be less than 45° and the turbine will be operating in a hybrid-mode, with a centrifugal-mode upstream and axial-mode downstream, harvesting energy from the diverted flow. This will result in a turbine having boomerang-shaped blades, as the green part depicted below. [[File:Hybrid_turbine.png]] ==Bypass flow and core discharge flow mixing== ''work in progress'' The hybrid turbine will be the last part crossed by the flow before the mixing. The hot flow will be abruptly expanded in the mixing area, increasing the static pressure. What about the temperature? if it's an isentropic irreversible expansion the temperature slightly decreases, if it's an adiabatic irreversible expansion the temperature decreases, but here no work is done so it's more like an non-ideal adiabatic free expansion for which the temperature [https://en.wikipedia.org/wiki/Free_expansion may] slightly change (to be verified, it does not change for an ideal gas). The static pressure increase and the velocity reduction will allow the fan to operate in an even more efficient way. An exhaust diffuser may have to be designed to control the expansion and the mixing in order to adjust the flow direction and velocity for the fan intake. The mixing can be seen as a '''preheat''', like the usual reheat to lesser extent, but before the fan has done its compression work, still allowing the nozzle to accelerate further the exhaust gas and thus providing a higher mass flow rate (= higher thrust and specific power). ==Fan operation== ''work in progress'' Main question: is the fan much less efficient when it blows hot air than cold air? Is this efficiency overtaken by the efficiency increase in transonic regime versus subsonic? Thrust depends on the mass flow rate of expelled material. Higher temperature means that the density is lower, and thus that the mass flow rate should be lower too. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000-series aluminium alloys. ==Thermodynamic analysis== Turbine engines are assimilated to a [https://en.wikipedia.org/wiki/Brayton_cycle Brayton cycle]. It consists of 4 processes: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion in two steps - first made by the turbine, the extracted work is used to drive both the compressor and the payload. In the case of a turbofan the payload is the fan. Then, the nozzle finishes the expansion and exhausts gas at atmosphere pressure. # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. In our approach, the expansion has to be considered as two different steps because there is a heat exchange between the two. Also, the fan doesn't operate in an adiabatic process since there is a heat input made by the core engine exhaust gas mixing. The two cycles, the core engine cycle and the fan cycle are now combined. The new process are: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion - made by the turbine ## isobaric heat rejection - made by the flow mixing ## adiabatic compression - made by the fan # adiabatic expansion - made by the nozzle # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. [[File:Thermodynamic_diagrams.png|center|250px]] ==Example implementation: full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' bypass ratio (BPR) because since flows are mixed before the fan intake, there is no clear separation between flows of the fan and of the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is thus the ratio between the sum of the fan duct mass flow rate and the core's mass flow rate over the core's mass flow rate (M_fan + M_core) / M_core. Let's take an example turbofan engine with a 13cm fan and a 9cm core (and compressor) diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both side ((13²-9²)/9²), and the BPR is in fact 2.09 (13²/9²). A great advantage of our aft-fan engine design is that both the first stage(s) of the compressor and the fan can operate at '''transonic''' speeds. A blade or a fan is said having a [http://en.wikipedia.org/wiki/Transonic transonic] operation when the flow passing through it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, while their root is generally subsonic too. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound, which depends on the temperature of the air flow. A transonic operation allows higher compression ratios to be achieved, with a lower efficiency than subsonic operation as a drawback, around 5% less because of the drag induced by shock waves. As a result, our design allows both the compressor entry stage and the fan to operate at transonic speeds, while they have the same rotation speed (one spool shaft), at high altitudes. This is possible because of the temperature difference in air passing through both. The compressor will breathe fresh air, which can go down as low as -57°C in altitude. In this case, the sonic speed is 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. If we fix a 200°C rise of this air flow, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. Having a transonic speed operation allows higher the compression ratios. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor, it's not able to have a high overall compression ratio (CR). With a transonic-operating first stage, we hope to have at least 2.0 CR for it, instead of the maximum 1.6 CR in subsonic operation, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Higher efficiency of the fan will allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. [[File:Engine core and fan.jpg|300px]] [[File:Engine core and fan side.jpg|350px]] ''The above schematics may not reflect actual proportions and have transparency inconsistencies.'' That said, the engine's core is the green internal tube. The fan sketched inside it, at the front of the engine, is the compressor's first stage rotor. The stator and other stage are not represented, and will obviously be placed on the conic part behind it. The red part is the annular [[Turbofan:Combustors|combustor]], not easy to see clearly in these pictures. '''The turbine is not represented''' either, but an axial rotor should be placed right after the combustor, and the hybrid turbine will be milled in the shaft in the curved part after a second axial turbine guide vane. We can see that the discharge of the turbine is mixed with the inlet air of the aft-mounted fan. The fan duct is thus longer than in traditional front-mounted engines, but the nozzle can be closer to the fan, allowing to work on both core and fan flow. [[Category:Turbofan|Alternative design]] 78fb0a3f9e1cb1a2cca77039d1a69514ab7cc5db 315 314 2012-06-02T03:52:59Z Vincent 1 /* Bypass flow and core discharge flow mixing */ more on temperature again wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine acting as the power plant behind it. Their flows mix at the exhaust, inside the engine for low bypass ratio engines and outside, in atmosphere, for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a hot gas flow, which would eventually corrode or induce fatigue on it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, [https://en.wikipedia.org/wiki/Heinkel_HeS_1 the HeS 1] ([http://airandspace.si.edu/collections/artifact.cfm?id=A19810039000 picture]), the first that actually flew a plane. From these two facts, we propose a novel design for turbofans: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. * The hybrid turbine uses both axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its linear course, while extracting some of its energy for shaft rotation work. The centrifugal part is crossed by the flow from the centre to the outer, contrary to what has been done in the HeS 1. * The air breathed by the fan would be the mixed flow of the traditional fresh intake and the turbine discharge, providing higher temperature to the fan flow. The main advantage of this hot air is the increase in speed of sound and the ability to drive the fan at transonic regime on the same spool than the compressor operating at transonic regime too. Multi-spool engines have been created to allow such features in classical engines. The engine will inevitably be longer because room has to be made for the flow-mixing section. This will make the engine slightly heavier too. ==Hybrid turbine== ''work in progress'' ''The reasons why gas turbines employ axial-turbine stages should be listed here''. For turbojets this is obviously due to the fact that the hot gas stream at engine discharge is what gives thrust, so the path of the hot gas from the turbine to the nozzle should be kept as linear as possible. For turbofans, the core engine exhaust also plays a role in providing thrust, generally around 10% of the total engine's thrust. Diverting the hot gas flow results in heavy loss of kinetic energy. Passing through a centrifugal-turbine extracts lots of energy from the gas and does not allow further work to be done from it. It also means that a centrifugal-turbine has to cope with much higher temperatures than axial-turbines, which can be a real show-stopper since axial-turbines already have important issues with temperature. In our case, we need to divert the hot gas flow in order to have it mixed properly with the cold air from the bypass. The diversion thus should be less than 45° and the turbine will be operating in a hybrid-mode, with a centrifugal-mode upstream and axial-mode downstream, harvesting energy from the diverted flow. This will result in a turbine having boomerang-shaped blades, as the green part depicted below. [[File:Hybrid_turbine.png]] ==Bypass flow and core discharge flow mixing== ''work in progress'' The hybrid turbine will be the last part crossed by the flow before the mixing. The hot flow will be abruptly expanded in the mixing area, converting the dynamic pressure (velocity) to static pressure. '''What about the temperature?''' if it's an isentropic irreversible expansion the temperature slightly decreases, if it's an adiabatic irreversible expansion the temperature decreases, but here no work is done so it's more like an non-ideal adiabatic free expansion for which the temperature [https://en.wikipedia.org/wiki/Free_expansion may] slightly change (to be verified, it does not change for an ideal gas); however, in a steady flow, an adiabatic diffuser (if we treat the heat exchange later in the mixing) increases both pressure and temperature. The static pressure increase and the velocity reduction will allow the fan to operate in an even more efficient way. An exhaust diffuser may have to be designed to control the expansion and the mixing in order to adjust the flow direction and velocity for the fan intake. The mixing can be seen as a '''preheat''', like the usual reheat to lesser extent, but before the fan has done its compression work, still allowing the nozzle to accelerate further the exhaust gas and thus providing a higher mass flow rate (= higher thrust and specific power). ==Fan operation== ''work in progress'' Main question: is the fan much less efficient when it blows hot air than cold air? Is this efficiency overtaken by the efficiency increase in transonic regime versus subsonic? Thrust depends on the mass flow rate of expelled material. Higher temperature means that the density is lower, and thus that the mass flow rate should be lower too. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000-series aluminium alloys. ==Thermodynamic analysis== Turbine engines are assimilated to a [https://en.wikipedia.org/wiki/Brayton_cycle Brayton cycle]. It consists of 4 processes: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion in two steps - first made by the turbine, the extracted work is used to drive both the compressor and the payload. In the case of a turbofan the payload is the fan. Then, the nozzle finishes the expansion and exhausts gas at atmosphere pressure. # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. In our approach, the expansion has to be considered as two different steps because there is a heat exchange between the two. Also, the fan doesn't operate in an adiabatic process since there is a heat input made by the core engine exhaust gas mixing. The two cycles, the core engine cycle and the fan cycle are now combined. The new process are: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion - made by the turbine ## isobaric heat rejection - made by the flow mixing ## adiabatic compression - made by the fan # adiabatic expansion - made by the nozzle # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. [[File:Thermodynamic_diagrams.png|center|250px]] ==Example implementation: full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' bypass ratio (BPR) because since flows are mixed before the fan intake, there is no clear separation between flows of the fan and of the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is thus the ratio between the sum of the fan duct mass flow rate and the core's mass flow rate over the core's mass flow rate (M_fan + M_core) / M_core. Let's take an example turbofan engine with a 13cm fan and a 9cm core (and compressor) diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both side ((13²-9²)/9²), and the BPR is in fact 2.09 (13²/9²). A great advantage of our aft-fan engine design is that both the first stage(s) of the compressor and the fan can operate at '''transonic''' speeds. A blade or a fan is said having a [http://en.wikipedia.org/wiki/Transonic transonic] operation when the flow passing through it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, while their root is generally subsonic too. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound, which depends on the temperature of the air flow. A transonic operation allows higher compression ratios to be achieved, with a lower efficiency than subsonic operation as a drawback, around 5% less because of the drag induced by shock waves. As a result, our design allows both the compressor entry stage and the fan to operate at transonic speeds, while they have the same rotation speed (one spool shaft), at high altitudes. This is possible because of the temperature difference in air passing through both. The compressor will breathe fresh air, which can go down as low as -57°C in altitude. In this case, the sonic speed is 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. If we fix a 200°C rise of this air flow, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. Having a transonic speed operation allows higher the compression ratios. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor, it's not able to have a high overall compression ratio (CR). With a transonic-operating first stage, we hope to have at least 2.0 CR for it, instead of the maximum 1.6 CR in subsonic operation, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Higher efficiency of the fan will allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. [[File:Engine core and fan.jpg|300px]] [[File:Engine core and fan side.jpg|350px]] ''The above schematics may not reflect actual proportions and have transparency inconsistencies.'' That said, the engine's core is the green internal tube. The fan sketched inside it, at the front of the engine, is the compressor's first stage rotor. The stator and other stage are not represented, and will obviously be placed on the conic part behind it. The red part is the annular [[Turbofan:Combustors|combustor]], not easy to see clearly in these pictures. '''The turbine is not represented''' either, but an axial rotor should be placed right after the combustor, and the hybrid turbine will be milled in the shaft in the curved part after a second axial turbine guide vane. We can see that the discharge of the turbine is mixed with the inlet air of the aft-mounted fan. The fan duct is thus longer than in traditional front-mounted engines, but the nozzle can be closer to the fan, allowing to work on both core and fan flow. [[Category:Turbofan|Alternative design]] 826e73f5ab17ab9e89df754c7aa361dbafd79965 318 315 2012-08-10T15:32:08Z Vincent 1 /* Example implementation: full transonic engine design in a single spool with 2.1 BPR */ few fixes wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine acting as the power plant behind it. Their flows mix at the exhaust, inside the engine for low bypass ratio engines and outside, in atmosphere, for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a hot gas flow, which would eventually corrode or induce fatigue on it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, [https://en.wikipedia.org/wiki/Heinkel_HeS_1 the HeS 1] ([http://airandspace.si.edu/collections/artifact.cfm?id=A19810039000 picture]), the first that actually flew a plane. From these two facts, we propose a novel design for turbofans: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. * The hybrid turbine uses both axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its linear course, while extracting some of its energy for shaft rotation work. The centrifugal part is crossed by the flow from the centre to the outer, contrary to what has been done in the HeS 1. * The air breathed by the fan would be the mixed flow of the traditional fresh intake and the turbine discharge, providing higher temperature to the fan flow. The main advantage of this hot air is the increase in speed of sound and the ability to drive the fan at transonic regime on the same spool than the compressor operating at transonic regime too. Multi-spool engines have been created to allow such features in classical engines. The engine will inevitably be longer because room has to be made for the flow-mixing section. This will make the engine slightly heavier too. ==Hybrid turbine== ''work in progress'' ''The reasons why gas turbines employ axial-turbine stages should be listed here''. For turbojets this is obviously due to the fact that the hot gas stream at engine discharge is what gives thrust, so the path of the hot gas from the turbine to the nozzle should be kept as linear as possible. For turbofans, the core engine exhaust also plays a role in providing thrust, generally around 10% of the total engine's thrust. Diverting the hot gas flow results in heavy loss of kinetic energy. Passing through a centrifugal-turbine extracts lots of energy from the gas and does not allow further work to be done from it. It also means that a centrifugal-turbine has to cope with much higher temperatures than axial-turbines, which can be a real show-stopper since axial-turbines already have important issues with temperature. In our case, we need to divert the hot gas flow in order to have it mixed properly with the cold air from the bypass. The diversion thus should be less than 45° and the turbine will be operating in a hybrid-mode, with a centrifugal-mode upstream and axial-mode downstream, harvesting energy from the diverted flow. This will result in a turbine having boomerang-shaped blades, as the green part depicted below. [[File:Hybrid_turbine.png]] ==Bypass flow and core discharge flow mixing== ''work in progress'' The hybrid turbine will be the last part crossed by the flow before the mixing. The hot flow will be abruptly expanded in the mixing area, converting the dynamic pressure (velocity) to static pressure. '''What about the temperature?''' if it's an isentropic irreversible expansion the temperature slightly decreases, if it's an adiabatic irreversible expansion the temperature decreases, but here no work is done so it's more like an non-ideal adiabatic free expansion for which the temperature [https://en.wikipedia.org/wiki/Free_expansion may] slightly change (to be verified, it does not change for an ideal gas); however, in a steady flow, an adiabatic diffuser (if we treat the heat exchange later in the mixing) increases both pressure and temperature. The static pressure increase and the velocity reduction will allow the fan to operate in an even more efficient way. An exhaust diffuser may have to be designed to control the expansion and the mixing in order to adjust the flow direction and velocity for the fan intake. The mixing can be seen as a '''preheat''', like the usual reheat to lesser extent, but before the fan has done its compression work, still allowing the nozzle to accelerate further the exhaust gas and thus providing a higher mass flow rate (= higher thrust and specific power). ==Fan operation== ''work in progress'' Main question: is the fan much less efficient when it blows hot air than cold air? Is this efficiency overtaken by the efficiency increase in transonic regime versus subsonic? Thrust depends on the mass flow rate of expelled material. Higher temperature means that the density is lower, and thus that the mass flow rate should be lower too. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000-series aluminium alloys. ==Thermodynamic analysis== Turbine engines are assimilated to a [https://en.wikipedia.org/wiki/Brayton_cycle Brayton cycle]. It consists of 4 processes: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion in two steps - first made by the turbine, the extracted work is used to drive both the compressor and the payload. In the case of a turbofan the payload is the fan. Then, the nozzle finishes the expansion and exhausts gas at atmosphere pressure. # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. In our approach, the expansion has to be considered as two different steps because there is a heat exchange between the two. Also, the fan doesn't operate in an adiabatic process since there is a heat input made by the core engine exhaust gas mixing. The two cycles, the core engine cycle and the fan cycle are now combined. The new process are: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion - made by the turbine ## isobaric heat rejection - made by the flow mixing ## adiabatic compression - made by the fan # adiabatic expansion - made by the nozzle # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. [[File:Thermodynamic_diagrams.png|center|250px]] ==Example implementation: full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' bypass ratio (BPR) because as flows are mixed before the fan intake, there is no clear separation between flows of the fan and of the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is thus the ratio between the sum of the fan duct mass flow rate and the core's mass flow rate over the core's mass flow rate (MFR_fan + MFR_core) / MFR_core. Let's take an example turbofan engine with a 13cm fan and a 9cm core diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both intakes ((13²-9²)/9²), and the BPR is in fact 2.09 (13²/9²). The main advantage of our aft-fan engine design is that both the first stage of the compressor and the fan can operate at '''transonic''' speeds. A blade or a fan is said having a [https://en.wikipedia.org/wiki/Transonic transonic] operation when the flow passing through the fan it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, their root staying generally subsonic. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound M, which depends on the temperature of the air flow. Transonic operation enables higher compression ratios than subsonic operation, with a lower efficiency as a drawback, around 5% less because of the drag induced by shock waves. It is likely that more noise is generated too, but it's not a concern for us, contrary to usual aircraft engines. The key to our dual-transonic operation mode is the temperature, which changes M. In high altitudes, or in particularly cold days, the temperature difference between the cold air entering the compressor and the hot mixed gas at the fan intake is large enough to enable the dual-transonic operation. The fresh air can go down as low as -57°C in altitude, which gives a sonic speed of 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. If we fix a 200°C rise of this air flow, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor for cost reasons, it's not able to have a high overall compression ratio (CR). With a transonic-operating first stage, we hope to have at least 2.0 CR for it, instead of the maximum 1.6 CR in subsonic operation, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Higher efficiency of the fan will allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. [[File:Engine core and fan.jpg|300px]] [[File:Engine core and fan side.jpg|350px]] ''The above schematics may not reflect actual proportions and have transparency inconsistencies.'' That said, the engine's core is the green internal tube. The fan sketched inside it, at the front of the engine, is the compressor's first stage rotor. The stator and other stage are not represented, and will obviously be placed on the conic part behind it. The red part is the annular [[Turbofan:Combustors|combustor]], not easy to see clearly in these pictures. '''The turbine is not represented''' either, but an axial rotor should be placed right after the combustor, and the hybrid turbine will be milled in the shaft in the curved part after a second axial turbine guide vane. We can see that the discharge of the turbine is mixed with the inlet air of the aft-mounted fan. The fan duct is thus longer than in traditional front-mounted engines, but the nozzle can be closer to the fan, allowing to work on both core and fan flow. [[Category:Turbofan|Alternative design]] 94114e4d42b360e76e216d8c439345132a0af952 319 318 2012-08-10T16:52:41Z Vincent 1 /* Fan operation */ about density wikitext text/x-wiki =Alternative design for turbofans= All recent jet engines have the same basic architecture: a fan at the front, the turbine engine acting as the power plant behind it. Their flows mix at the exhaust, inside the engine for low bypass ratio engines and outside, in atmosphere, for high bypass ratio engines. Early designs of turbofans were actually created by adding a ducted fan on the aft part of existing turbojet engines. They were not so bad in terms of efficiency compared to front-mounting engines, but the main issue was that the fan was evolving in a hot gas flow, which would eventually corrode or induce fatigue on it more than when it blows fresh air as in a front-mounted fan design. A second fact is that high-efficiency engines, or modern engines, all use axial-type compressor and axial-type turbine. Some early or less power-requiring designs feature a centrifugal-type compressor, and only one engine to our knowledge had a centrifugal turbine, [https://en.wikipedia.org/wiki/Heinkel_HeS_1 the HeS 1] ([http://airandspace.si.edu/collections/artifact.cfm?id=A19810039000 picture]), the first that actually flew a plane. From these two facts, we propose a novel design for turbofans: '''an axial-compressor, hybrid-turbine, aft-mounted ducted fan'''. * The hybrid turbine uses both axial and centrifugal designs, in which the hot gas flow would be slightly diverted from its linear course, while extracting some of its energy for shaft rotation work. The centrifugal part is crossed by the flow from the centre to the outer, contrary to what has been done in the HeS 1. * The air breathed by the fan would be the mixed flow of the traditional fresh intake and the turbine discharge, providing higher temperature to the fan flow. The main advantage of this hot air is the increase in speed of sound and the ability to drive the fan at transonic regime on the same spool than the compressor operating at transonic regime too. Multi-spool engines have been created to allow such features in classical engines. The engine will inevitably be longer because room has to be made for the flow-mixing section. This will make the engine slightly heavier too. ==Hybrid turbine== ''work in progress'' ''The reasons why gas turbines employ axial-turbine stages should be listed here''. For turbojets this is obviously due to the fact that the hot gas stream at engine discharge is what gives thrust, so the path of the hot gas from the turbine to the nozzle should be kept as linear as possible. For turbofans, the core engine exhaust also plays a role in providing thrust, generally around 10% of the total engine's thrust. Diverting the hot gas flow results in heavy loss of kinetic energy. Passing through a centrifugal-turbine extracts lots of energy from the gas and does not allow further work to be done from it. It also means that a centrifugal-turbine has to cope with much higher temperatures than axial-turbines, which can be a real show-stopper since axial-turbines already have important issues with temperature. In our case, we need to divert the hot gas flow in order to have it mixed properly with the cold air from the bypass. The diversion thus should be less than 45° and the turbine will be operating in a hybrid-mode, with a centrifugal-mode upstream and axial-mode downstream, harvesting energy from the diverted flow. This will result in a turbine having boomerang-shaped blades, as the green part depicted below. [[File:Hybrid_turbine.png]] ==Bypass flow and core discharge flow mixing== ''work in progress'' The hybrid turbine will be the last part crossed by the flow before the mixing. The hot flow will be abruptly expanded in the mixing area, converting the dynamic pressure (velocity) to static pressure. '''What about the temperature?''' if it's an isentropic irreversible expansion the temperature slightly decreases, if it's an adiabatic irreversible expansion the temperature decreases, but here no work is done so it's more like an non-ideal adiabatic free expansion for which the temperature [https://en.wikipedia.org/wiki/Free_expansion may] slightly change (to be verified, it does not change for an ideal gas); however, in a steady flow, an adiabatic diffuser (if we treat the heat exchange later in the mixing) increases both pressure and temperature. The static pressure increase and the velocity reduction will allow the fan to operate in an even more efficient way. An exhaust diffuser may have to be designed to control the expansion and the mixing in order to adjust the flow direction and velocity for the fan intake. The mixing can be seen as a '''preheat''', like the usual reheat to lesser extent, but before the fan has done its compression work, still allowing the nozzle to accelerate further the exhaust gas and thus providing a higher mass flow rate (= higher thrust and specific power). ==Fan operation== ''work in progress'' Main question: is the fan much less efficient when it blows hot air than cold air? Is this efficiency overtaken by the efficiency increase in transonic regime versus subsonic? The air density is 1.9 times lower at 150°C (423K) compared to -50°C (223K), which would mean that the fan is nearly two times less efficient, if the pressure stays constant during the heating, but does it? Probably not, since the hot gas is expanded at core's discharge, it trades its dynamic pressure to higher static pressure. Thrust depends on the mass flow rate of expelled material. Higher temperature means that the density is lower, and thus that the mass flow rate should be lower too. Properly mixing the two flows would allow the fan to be build with metals supporting low temperatures, like 2000- or 7000-series aluminium alloys. ==Thermodynamic analysis== Turbine engines are assimilated to a [https://en.wikipedia.org/wiki/Brayton_cycle Brayton cycle]. It consists of 4 processes: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion in two steps - first made by the turbine, the extracted work is used to drive both the compressor and the payload. In the case of a turbofan the payload is the fan. Then, the nozzle finishes the expansion and exhausts gas at atmosphere pressure. # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. In our approach, the expansion has to be considered as two different steps because there is a heat exchange between the two. Also, the fan doesn't operate in an adiabatic process since there is a heat input made by the core engine exhaust gas mixing. The two cycles, the core engine cycle and the fan cycle are now combined. The new process are: # adiabatic compression - made by the compressor # isobaric heat addition - made by the combustion # adiabatic expansion - made by the turbine ## isobaric heat rejection - made by the flow mixing ## adiabatic compression - made by the fan # adiabatic expansion - made by the nozzle # isobaric heat rejection - made when the exhaust gas returns in the atmosphere. [[File:Thermodynamic_diagrams.png|center|250px]] ==Example implementation: full transonic engine design in a single spool with 2.1 BPR== We speak here of ''virtual'' bypass ratio (BPR) because as flows are mixed before the fan intake, there is no clear separation between flows of the fan and of the engine's core. However, there is still an inlet area for the compressor and one for the fan, and the ratio between the two mass flow rates going into each is what we call the virtual bypass ratio (VBPR). The real bypass ratio (BPR) is thus the ratio between the sum of the fan duct mass flow rate and the core's mass flow rate over the core's mass flow rate (MFR_fan + MFR_core) / MFR_core. Let's take an example turbofan engine with a 13cm fan and a 9cm core diameter. The VBPR for this engine is around 1.08 if we consider the inlet flow speeds to be identical on both intakes ((13²-9²)/9²), and the BPR is in fact 2.09 (13²/9²). The main advantage of our aft-fan engine design is that both the first stage of the compressor and the fan can operate at '''transonic''' speeds. A blade or a fan is said having a [https://en.wikipedia.org/wiki/Transonic transonic] operation when the flow passing through the fan it is subsonic, but its rotation speed makes the blades' tips move at supersonic velocities, their root staying generally subsonic. It thus depends on three main factors: the diameter of the fan on which the blade is mounted, the rotation speed of the fan, and the speed of sound M, which depends on the temperature of the air flow. Transonic operation enables higher compression ratios than subsonic operation, with a lower efficiency as a drawback, around 5% less because of the drag induced by shock waves. It is likely that more noise is generated too, but it's not a concern for us, contrary to usual aircraft engines. The key to our dual-transonic operation mode is the temperature, which changes M. In high altitudes, or in particularly cold days, the temperature difference between the cold air entering the compressor and the hot mixed gas at the fan intake is large enough to enable the dual-transonic operation. The fresh air can go down as low as -57°C in altitude, which gives a sonic speed of 295m/s. On the other hand, the air that the fan blows will be preheated by the turbine exhaust gas. If we fix a 200°C rise of this air flow, the sonic speed can be around 430m/s. If we take back our example above with the 13cm fan and 9cm compressor with the 200°C heating of the fan inlet flow, '''we achieve sonic speed''' with the same rotational speed, more than 60000rpm, '''for both the compressor first stage and the fan, which is unique for a 2.1 or even a 1.1 BPR turbofan engine'''. The reason why we try to have a transonic operation on the compressor, even more than on the fan, is that since we aim a low number of stages for the compressor for cost reasons, it's not able to have a high overall compression ratio (CR). With a transonic-operating first stage, we hope to have at least 2.0 CR for it, instead of the maximum 1.6 CR in subsonic operation, which would greatly improve the overall CR of the compressor. Depending on the temperature rise induced by this first stage, the second stage may be able to operate at transonic or near sonic speeds too, although unlikely. Higher efficiency of the fan will allow higher mass flow rate and thus higher thrust of the engine, which is obviously great too. [[File:Engine core and fan.jpg|300px]] [[File:Engine core and fan side.jpg|350px]] ''The above schematics may not reflect actual proportions and have transparency inconsistencies.'' That said, the engine's core is the green internal tube. The fan sketched inside it, at the front of the engine, is the compressor's first stage rotor. The stator and other stage are not represented, and will obviously be placed on the conic part behind it. The red part is the annular [[Turbofan:Combustors|combustor]], not easy to see clearly in these pictures. '''The turbine is not represented''' either, but an axial rotor should be placed right after the combustor, and the hybrid turbine will be milled in the shaft in the curved part after a second axial turbine guide vane. We can see that the discharge of the turbine is mixed with the inlet air of the aft-mounted fan. The fan duct is thus longer than in traditional front-mounted engines, but the nozzle can be closer to the fan, allowing to work on both core and fan flow. [[Category:Turbofan|Alternative design]] a8034f8a4d3091f83958485c0514a0690a82d4ca 0 16 307 271 2012-05-28T23:10:55Z Vincent 1 adding ssto blog wikitext text/x-wiki =Resources= This page gathers available documentation on the numerous subjects linked the project, including rocket science, turbofans, aerodynamics, amateur rocketry and SSTO. An [[Aero formulas|internal page]] is being created containing useful formulas related to thermodynamics and aerodynamics. ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/l amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * An important nasaspaceflight [http://forum.nasaspaceflight.com/index.php?topic=2847.0 forum thread] where ideas on micro-rocketry to orbit are discussed and shared. * A kind of spin-off of the above thread is the [http://orbitalaspirations.blogspot.com/ orbital aspirations weblog]. It was recently created by Ed LeBouthillier and is already filled with lots of information on scaling down rocket equations to micro-rocketry, SSTO, reference papers, news of the domain and so on. * Another blog is dedicated to SSTO: [http://exoscientist.blogspot.fr/ Polymath/exoscientist]. * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] by Klaus Hünecke, 1997. * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ===Other Aerodynamics=== * [http://www.desktop.aero/appliedaero/preface/welcome.html Applied Aerodynamics: A Digital Textbook] by Desktop Aeronautics, Inc. 2007. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. fc4a8eb4429e0aed36f2db3349b9e60f82dfe78f 308 307 2012-05-29T00:33:14Z Vincent 1 /* Web pages */ fixing link wikitext text/x-wiki =Resources= This page gathers available documentation on the numerous subjects linked the project, including rocket science, turbofans, aerodynamics, amateur rocketry and SSTO. An [[Aero formulas|internal page]] is being created containing useful formulas related to thermodynamics and aerodynamics. ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/ amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * An important nasaspaceflight [http://forum.nasaspaceflight.com/index.php?topic=2847.0 forum thread] where ideas on micro-rocketry to orbit are discussed and shared. * A kind of spin-off of the above thread is the [http://orbitalaspirations.blogspot.com/ orbital aspirations weblog]. It was recently created by Ed LeBouthillier and is already filled with lots of information on scaling down rocket equations to micro-rocketry, SSTO, reference papers, news of the domain and so on. * Another blog is dedicated to SSTO: [http://exoscientist.blogspot.fr/ Polymath/exoscientist]. * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] by Klaus Hünecke, 1997. * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ===Other Aerodynamics=== * [http://www.desktop.aero/appliedaero/preface/welcome.html Applied Aerodynamics: A Digital Textbook] by Desktop Aeronautics, Inc. 2007. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. 0aa3f611538744b8cde91c238ba61434c6737ce6 316 308 2012-06-02T04:12:39Z Vincent 1 /* Books (online links) */ thermodynamics book, shanthini, 2006 wikitext text/x-wiki =Resources= This page gathers available documentation on the numerous subjects linked the project, including rocket science, turbofans, aerodynamics, amateur rocketry and SSTO. An [[Aero formulas|internal page]] is being created containing useful formulas related to thermodynamics and aerodynamics. ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/ amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * An important nasaspaceflight [http://forum.nasaspaceflight.com/index.php?topic=2847.0 forum thread] where ideas on micro-rocketry to orbit are discussed and shared. * A kind of spin-off of the above thread is the [http://orbitalaspirations.blogspot.com/ orbital aspirations weblog]. It was recently created by Ed LeBouthillier and is already filled with lots of information on scaling down rocket equations to micro-rocketry, SSTO, reference papers, news of the domain and so on. * Another blog is dedicated to SSTO: [http://exoscientist.blogspot.fr/ Polymath/exoscientist]. * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] by Klaus Hünecke, 1997. * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ===Other: Thermodynamics, Aerodynamics, general physics=== * [http://www.desktop.aero/appliedaero/preface/welcome.html Applied Aerodynamics: A Digital Textbook] by Desktop Aeronautics, Inc. 2007. * [http://www.rshanthini.com/ThermoBook.htm Thermodynamics for Beginners, with worked examples] by R. Shanthini. 2006. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. fabefc1209081a6047ef981e0ff9a93bde0bd0b5 330 316 2012-08-16T01:47:46Z Vincent 1 /* Web pages */ +selenian boondocks wikitext text/x-wiki =Resources= This page gathers available documentation on the numerous subjects linked the project, including rocket science, turbofans, aerodynamics, amateur rocketry and SSTO. An [[Aero formulas|internal page]] is being created containing useful formulas related to thermodynamics and aerodynamics. ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/ amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * An important nasaspaceflight [http://forum.nasaspaceflight.com/index.php?topic=2847.0 forum thread] where ideas on micro-rocketry to orbit are discussed and shared. * A kind of spin-off of the above thread is the [http://orbitalaspirations.blogspot.com/ orbital aspirations weblog]. It was recently created by Ed LeBouthillier and is already filled with lots of information on scaling down rocket equations to micro-rocketry, SSTO, reference papers, news of the domain and so on. * Another blog talking about SSTO and air-to-orbit: [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html selenian boondocks]. * Another blog is dedicated to SSTO: [http://exoscientist.blogspot.fr/ Polymath/exoscientist]. * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] by Klaus Hünecke, 1997. * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ===Other: Thermodynamics, Aerodynamics, general physics=== * [http://www.desktop.aero/appliedaero/preface/welcome.html Applied Aerodynamics: A Digital Textbook] by Desktop Aeronautics, Inc. 2007. * [http://www.rshanthini.com/ThermoBook.htm Thermodynamics for Beginners, with worked examples] by R. Shanthini. 2006. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. b13fb3ca1810170c4f33e530cf1f95f016eed05e 0 56 309 296 2012-05-31T03:36:56Z Vincent 1 adding another p variable: the momentum wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] or [https://en.wikipedia.org/wiki/Heat_capacity_ratio Heat capacity ratio] (adiabatic process in thermodynamics) | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | ρ (rho) | [https://en.wikipedia.org/wiki/Density Density] | kg.m^-3 (kg per cubic meter) |- | C, C_p, C_V | [https://en.wikipedia.org/wiki/Heat_capacity#Metrology Heat capacity], general, at constant pressure, at constant volume. | J.K^-1 (Joule per Kelvin) |- | G | [https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] | J (Joule) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: total energy of a thermodynamic system. | J (Joule) |- | ΔH_vap or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | M | [https://en.wikipedia.org/wiki/Mach_number Mach number] | no unit |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature. T₀ or T_t is the [https://en.wikipedia.org/wiki/Stagnation_temperature stagnation temperature]. | K (Kelvin) |- | S | [https://en.wikipedia.org/wiki/Entropy Entropy] | J.K^-1 (Joule per Kelvin) |- | U | [https://en.wikipedia.org/wiki/Internal_energy Internal energy] of a system (see first law of Thermodynamics below) | J (Joule) |- | V | Volume | m³ (cubic meter) |- | W | [https://en.wikipedia.org/wiki/Work_(thermodynamics) Work]: mechanical constraints on the system. | J (Joule) |- | a | [https://en.wikipedia.org/wiki/Speed_of_sound Speed of sound] in medium (used to calculate Mach number) | m.s^-1 |- | c | Velocity of a flow in thermodynamics, also noted V; generally noted u in fluid dynamics. | m.s^-1 |- | n | Quantity of matter | mol (mole) |- | p | Pressure. p_t is the [https://en.wikipedia.org/wiki/Stagnation_pressure stagnation pressure]. | Pa (Pascal) |- | p | [https://en.wikipedia.org/wiki/Momentum Momentum] p = m*v, with m the mass and v the velocity, not to be confused with volume. | kg.m.s^-1 |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | G | [https://en.wikipedia.org/wiki/Gravitational_constant Gravitational constant] | 6.674 | m³.kg^-1.s^-2 |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |style="background:white"| {{SERVER}}/images/formulas_mirror/pvnrtk.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/clausius-clapeyron.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/QeqmL.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. Delta H equals Q only when pressure is constant (isobaric). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dUeqdQmindW.png |[https://en.wikipedia.org/wiki/First_law_of_thermodynamics First law of thermodynamics] |Variations of internal energy of a system between two states is the sum of the received heat and work (minus the ''given'' work). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/enthalpy.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] |Total amount of energy of a system, defined as the sum of the internal energy U of the system and pressure * volume at the boundary of the system and its environment. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/workExpand.png |Work of gas expansion. |Work done by expanding an ideal gas. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/entropy_dueqtdsmpdv.png |[https://en.wikipedia.org/wiki/Internal_energy Internal energy] change related to [https://en.wikipedia.org/wiki/Entropy entropy] |Internal energy related to entropy variation for a closed system in thermal equilibrium ([https://en.wikipedia.org/wiki/Fundamental_thermodynamic_relation fundamental thermodynamic relation]). |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dheqtds.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] change |Enthalpy change depending on entropy and pressure changes, equation created from the mix of the basic ones above. |- |ΔS_universe = ΔS_surroundings + ΔS_system |Entropy variation as a whole. |Entropy variation of a system is generally compensated by the inverse variation of the surroundings, not including losses. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/dS.png |[https://en.wikipedia.org/wiki/Second_law_of_thermodynamics Second law of thermodynamics] |A change in the entropy of a system is the infinitesimal transfer of heat to a closed system driving a reversible process, divided by the equilibrium temperature of the system. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/gibbs.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] / Free enthalpy |Useful work obtainable from a system at '''isobaric''' and '''isothermal''' conditions. Since H is U + pV, it can be replaced in the equation, making G = H - TS. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/deltaG.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] variation. |If ΔG < 0, the system's transformation can be spontaneous, if ΔG = 0 the transformation is inversible and the system is in an equilibrium state, if ΔG > 0 it can't be spontaneous. |- |style="background:white"| {{SERVER}}/images/formulas_mirror/density_ideal.png |[https://en.wikipedia.org/wiki/Density#Changes_of_density Density] of an ideal gas. |M is molar mass. This means that the density of an ideal gas can be doubled by doubling the pressure, or by halving the absolute temperature. |} d23f45dc05391ee2b9a374720d1551be0466a3db 347 309 2012-11-01T01:22:48Z Vincent 1 /* List of equations */ fixed formulas images colors wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] or [https://en.wikipedia.org/wiki/Heat_capacity_ratio Heat capacity ratio] (adiabatic process in thermodynamics) | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | ρ (rho) | [https://en.wikipedia.org/wiki/Density Density] | kg.m^-3 (kg per cubic meter) |- | C, C_p, C_V | [https://en.wikipedia.org/wiki/Heat_capacity#Metrology Heat capacity], general, at constant pressure, at constant volume. | J.K^-1 (Joule per Kelvin) |- | G | [https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] | J (Joule) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: total energy of a thermodynamic system. | J (Joule) |- | ΔH_vap or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | M | [https://en.wikipedia.org/wiki/Mach_number Mach number] | no unit |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature. T₀ or T_t is the [https://en.wikipedia.org/wiki/Stagnation_temperature stagnation temperature]. | K (Kelvin) |- | S | [https://en.wikipedia.org/wiki/Entropy Entropy] | J.K^-1 (Joule per Kelvin) |- | U | [https://en.wikipedia.org/wiki/Internal_energy Internal energy] of a system (see first law of Thermodynamics below) | J (Joule) |- | V | Volume | m³ (cubic meter) |- | W | [https://en.wikipedia.org/wiki/Work_(thermodynamics) Work]: mechanical constraints on the system. | J (Joule) |- | a | [https://en.wikipedia.org/wiki/Speed_of_sound Speed of sound] in medium (used to calculate Mach number) | m.s^-1 |- | c | Velocity of a flow in thermodynamics, also noted V; generally noted u in fluid dynamics. | m.s^-1 |- | n | Quantity of matter | mol (mole) |- | p | Pressure. p_t is the [https://en.wikipedia.org/wiki/Stagnation_pressure stagnation pressure]. | Pa (Pascal) |- | p | [https://en.wikipedia.org/wiki/Momentum Momentum] p = m*v, with m the mass and v the velocity, not to be confused with volume. | kg.m.s^-1 |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | G | [https://en.wikipedia.org/wiki/Gravitational_constant Gravitational constant] | 6.674 | m³.kg^-1.s^-2 |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |{{SERVER}}/images/formulas_mirror/pvnrtk_neg.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |{{SERVER}}/images/formulas_mirror/clausius-clapeyron_neg.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |{{SERVER}}/images/formulas_mirror/QeqmL_neg.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. Delta H equals Q only when pressure is constant (isobaric). |- |{{SERVER}}/images/formulas_mirror/dUeqdQmindW_neg.png |[https://en.wikipedia.org/wiki/First_law_of_thermodynamics First law of thermodynamics] |Variations of internal energy of a system between two states is the sum of the received heat and work (minus the ''given'' work). |- |{{SERVER}}/images/formulas_mirror/enthalpy_neg.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] |Total amount of energy of a system, defined as the sum of the internal energy U of the system and pressure * volume at the boundary of the system and its environment. |- |{{SERVER}}/images/formulas_mirror/workExpand_neg.png |Work of gas expansion. |Work done by expanding an ideal gas. |- |{{SERVER}}/images/formulas_mirror/entropy_dueqtdsmpdv_neg.png |[https://en.wikipedia.org/wiki/Internal_energy Internal energy] change related to [https://en.wikipedia.org/wiki/Entropy entropy] |Internal energy related to entropy variation for a closed system in thermal equilibrium ([https://en.wikipedia.org/wiki/Fundamental_thermodynamic_relation fundamental thermodynamic relation]). |- |{{SERVER}}/images/formulas_mirror/dheqtds_neg.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] change |Enthalpy change depending on entropy and pressure changes, equation created from the mix of the basic ones above. |- |ΔS_universe = ΔS_surroundings + ΔS_system |Entropy variation as a whole. |Entropy variation of a system is generally compensated by the inverse variation of the surroundings, not including losses. |- |{{SERVER}}/images/formulas_mirror/dS_neg.png |[https://en.wikipedia.org/wiki/Second_law_of_thermodynamics Second law of thermodynamics] |A change in the entropy of a system is the infinitesimal transfer of heat to a closed system driving a reversible process, divided by the equilibrium temperature of the system. |- |{{SERVER}}/images/formulas_mirror/gibbs_neg.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] / Free enthalpy |Useful work obtainable from a system at '''isobaric''' and '''isothermal''' conditions. Since H is U + pV, it can be replaced in the equation, making G = H - TS. |- |{{SERVER}}/images/formulas_mirror/deltaG_neg.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] variation. |If ΔG < 0, the system's transformation can be spontaneous, if ΔG = 0 the transformation is inversible and the system is in an equilibrium state, if ΔG > 0 it can't be spontaneous. |- |{{SERVER}}/images/formulas_mirror/density_ideal_neg.png |[https://en.wikipedia.org/wiki/Density#Changes_of_density Density] of an ideal gas. |M is molar mass. This means that the density of an ideal gas can be doubled by doubling the pressure, or by halving the absolute temperature. |} f4c92f9a3ac43aae0011efc085006b1f66303aae 0 9 317 288 2012-08-02T11:09:23Z Vincent 1 /* Sensors */ pressure wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [https://www.armadeus.com/ Armadeus], based on an ARM (FreeScale) processor. Moreover, it [https://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [https://xenomai.org/ Xenomai] Linux-based RTOS. Armadeus board integration has a [[Armadeus|dedicated page]]. Other interesting embedded computer boards: the [https://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1] and the [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20]. They do not include a FPGA but are cheaper, the first has a great temperature range, and the second is more powerful and smaller. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [https://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. ===Telemetry=== It seems that the 900MHz version of the ZigBee communication standard is able to transmit at around 100kbps up to 10km. Taken from th ArduPilot page: Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [https://www.adafruit.com/products/126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [https://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [https://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: quite common nowadays, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the theoretically computed strength the mechanical structure has to sustain. To chose a sensor, sparkfun wrote an [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and might be replaced by a camera sensor for low rotation rates. * Magnetometer (3D compass): lots of sensors exist too, for example the [https://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag] and [https://www.sparkfun.com/products/8128 SCP1000], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * Static pressure sensor, like [https://www.freescale.com/files/sensors/doc/data_sheet/MP3H6115A.pdf Freescale's MP3H6115A] and another for the extended range of high-altitude. * Thermometer: for systems health monitoring, like engines temperature. * GPS if USAF and sensors allow it in flight altitude. * Camera: 8-bit data port if possible, like the TCM8230MD sensor. Some ARM processors (i.MX) feature the Camera/CMOS Sensor Interface (CSI) and hardware-accelerated processing or compression from this port. Horizon sensor is provided by a camera. * Pitot tubes even exist in stores (like [https://store.diydrones.com/Kit_MPXV7002DP_p/kt-mpxv7002dp-01.htm DIYDrones])! * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [https://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. 15d870a27c07e88ba152cc7b38d8f788eae45967 320 317 2012-08-11T02:51:04Z Vincent 1 /* Sensors */ static pressure ports wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [https://www.armadeus.com/ Armadeus], based on an ARM (FreeScale) processor. Moreover, it [https://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [https://xenomai.org/ Xenomai] Linux-based RTOS. Armadeus board integration has a [[Armadeus|dedicated page]]. Other interesting embedded computer boards: the [https://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1] and the [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20]. They do not include a FPGA but are cheaper, the first has a great temperature range, and the second is more powerful and smaller. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [https://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. ===Telemetry=== It seems that the 900MHz version of the ZigBee communication standard is able to transmit at around 100kbps up to 10km. Taken from th ArduPilot page: Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [https://www.adafruit.com/products/126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [https://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [https://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: quite common nowadays, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the theoretically computed strength the mechanical structure has to sustain. To chose a sensor, sparkfun wrote an [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and might be replaced by a camera sensor for low rotation rates. * Magnetometer (3D compass): lots of sensors exist too, for example the [https://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag] and [https://www.sparkfun.com/products/8128 SCP1000], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * Static pressure sensor, like [https://www.freescale.com/files/sensors/doc/data_sheet/MP3H6115A.pdf Freescale's MP3H6115A] and another for the extended range of high-altitude. The pressure sensor requires vents, or static ports, in the fuselage to exchange outside air with inside pressure chamber. Details about the size and number of these ports are explained [http://www.adeptrocketry.com/A1ds.htm here] for model rocketry. * Thermometer: for systems health monitoring, like engines temperature. * GPS if USAF and sensors allow it in flight altitude. * Camera: 8-bit data port if possible, like the TCM8230MD sensor. Some ARM processors (i.MX) feature the Camera/CMOS Sensor Interface (CSI) and hardware-accelerated processing or compression from this port. Horizon sensor is provided by a camera. * Pitot tubes even exist in stores (like [https://store.diydrones.com/Kit_MPXV7002DP_p/kt-mpxv7002dp-01.htm DIYDrones])! * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [https://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. 241e7f5ca223f4074873dd7915876eca03d141a9 0 55 321 267 2012-08-11T19:14:15Z Vincent 1 mass flow rate and conservation of mass wikitext text/x-wiki =Flight at high altitude= Some information is summarized in the main page already, in the [[Main_Page#The_aircraft|aircraft section]]. ==Gas properties and altitude== [https://en.wikipedia.org/wiki/Density_of_air Density] is used to calculate [https://en.wikipedia.org/wiki/Lift_(force) lift] of an wing and [https://en.wikipedia.org/wiki/Thrust thrust] of an engine amongst other things. We absolutely need to know approximately what air densities will be faced in order to design everything. An atmosphere model should be used for design, but can also be used at runtime to verify that the actual atmosphere is within prediction range. Air density depends on pressure. [http://www.respirometry.org/look-up-table/barometric-pressure-vs-altitude This table] gives atmospheric pressure and temperature depending on altitude. We can see that a tenth of ground atmospheric pressure is met at around 16km altitude, and a hundredth of it at around 31km altitude. Air density in the atmosphere is also related to the ratio of air vapour in it, as indicated on [https://wahiduddin.net/calc/density_altitude.htm this page]. The page also contains lots of formulas and calculators, most importantly the ''air density calculator'' that we'll use right below. Water vapour however, is much more rare when temperature goes down, as it does in the higher troposphere or low to mid stratosphere that we're aiming. The calculator gives us, with temperature and pressure values taken from the table mentioned above, values for density of: * 1.214 kg/m^3 at sea level (15°C) * 0.1877 kg/m^3 at 15km altitude (-57°C) * 0.0441 kg/m^3 at 25km altitude (-52°C) * 0.017 kg/m^3 at 30km altitude (-46°C) ==Turbofan engine's Mass flow rate calculation== One way of calculating the MFR is to use the [https://en.wikipedia.org/wiki/Continuity_equation#Fluid_dynamics continuity equation]. The mass of gas leaving the engine is the same than the mass of gas entering the engine, for which we know the density, plus the mass of the fuel, which is much lower than the mass of air. It's the velocity difference between input and output that creates the thrust. ==Approaches overview== ===High engine power at low air density=== '''Is it possible to have a low total pressure ratio engine operating at subsonic inlet speeds and low air density?''' The MiG 25 has supersonic inlet, which allows him to have a significant pressure increase before the compressor actually gives energy to the flow. A subsonic input air flow in the high-altitude conditions is likely to not provide enough oxygen for the combustion to maintain by itself, or a too poor mass flow rate to the turbine. ===High lift at low air density=== High engine power in low air density generally means supersonic flight, or at least, high flight speeds, which in return increase the lift of the aircraft or decrease its wingspan. Our next step is to make some calculations of the required winged area for subsonic low-density air travel, and assess the feasibility of our air launch to orbit project. c85ea8541430756f68d7ff07e683c773d3e6e5c9 340 321 2012-10-15T03:19:43Z Vincent 1 /* Gas properties and altitude */ typo wikitext text/x-wiki =Flight at high altitude= Some information is summarized in the main page already, in the [[Main_Page#The_aircraft|aircraft section]]. ==Gas properties and altitude== [https://en.wikipedia.org/wiki/Density_of_air Density] is used to calculate [https://en.wikipedia.org/wiki/Lift_(force) lift] of an wing and [https://en.wikipedia.org/wiki/Thrust thrust] of an engine amongst other things. We absolutely need to know approximately what air densities will be faced in order to design everything. An atmosphere model should be used for design, but can also be used at runtime to verify that the actual atmosphere is within prediction range. Air density depends on pressure. [http://www.respirometry.org/look-up-table/barometric-pressure-vs-altitude This table] gives atmospheric pressure and temperature depending on altitude. We can see that a tenth of ground atmospheric pressure (ground-level: 1atm) is met at around 16km altitude, and a hundredth of it at around 31km altitude. Air density in the atmosphere is also related to the ratio of water vapour in it, as indicated on [https://wahiduddin.net/calc/density_altitude.htm this page]. The page also contains lots of formulas and calculators, most importantly the ''air density calculator'' that we'll use right below. Water vapour however, is much more rare when temperature goes down, as it does in the higher troposphere or low to mid stratosphere that we're aiming. The calculator gives us, with temperature and pressure values taken from the table mentioned above, values for density of: * 1.214 kg/m^3 at sea level (15°C) * 0.1877 kg/m^3 at 15km altitude (-57°C) * 0.0441 kg/m^3 at 25km altitude (-52°C) * 0.017 kg/m^3 at 30km altitude (-46°C) ==Turbofan engine's Mass flow rate calculation== One way of calculating the MFR is to use the [https://en.wikipedia.org/wiki/Continuity_equation#Fluid_dynamics continuity equation]. The mass of gas leaving the engine is the same than the mass of gas entering the engine, for which we know the density, plus the mass of the fuel, which is much lower than the mass of air. It's the velocity difference between input and output that creates the thrust. ==Approaches overview== ===High engine power at low air density=== '''Is it possible to have a low total pressure ratio engine operating at subsonic inlet speeds and low air density?''' The MiG 25 has supersonic inlet, which allows him to have a significant pressure increase before the compressor actually gives energy to the flow. A subsonic input air flow in the high-altitude conditions is likely to not provide enough oxygen for the combustion to maintain by itself, or a too poor mass flow rate to the turbine. ===High lift at low air density=== High engine power in low air density generally means supersonic flight, or at least, high flight speeds, which in return increase the lift of the aircraft or decrease its wingspan. Our next step is to make some calculations of the required winged area for subsonic low-density air travel, and assess the feasibility of our air launch to orbit project. 2644735cba6708b9c65b6f27d52b6b6d1a4f28e2 0 61 322 2012-08-13T02:52:39Z Vincent 1 parameters to take into account to calculate an approximative rocket mass wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important points about rockets launched from balloon or from aircraft, mainly what minimum mass we can expect. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require depends on several factors: * rocket launch altitude, the higher it is, the less delta V is required to face atmospheric drag and to exit the atmosphere and actually start the useful horizontal delta V. The planned altitude is 30km ±5km. * rocket launch speed, which is close to the speed of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|high-altitude flight strategy]], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical component, likely consistent with the rocket flight path initial vector. ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that [deltaV image], where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta v doesn't take into account gravity and other forces acting on the vehicle. * the exhaust velocity of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: [v_e image]. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the best most efficient rocket engine, * thrust is related to this exhaust velocity in the following way [thrust image]. It is not directly related to the delta v, but it is related to the mass of the payload that has to be accelerated. Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the dry mass of the rocket, once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket, the more efficient it will be in term of delta V. To make a light rocket, high quality materials, thus expensive, must be used. What takes the most volume in a rocket is the fuel and oxidizer tanks. If tanks are solid enough, and this may be a requirements depending on what they contain, they can be used as part of the structure of the rocket. If they are not solid, a external skeleton has to be build that will undergo all mechanical forces that apply to the vehicle. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. * the mass of the engine itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The trust to mass ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. e1874aa6ba8c9b44ad08ab9026e42b1ead11e8c5 323 322 2012-08-13T03:26:59Z Vincent 1 images and orbital speed wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important points about rockets launched from balloon or from aircraft, mainly what minimum mass we can expect. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require depends on several factors: * rocket launch altitude, the higher it is, the less delta V is required to face atmospheric drag and to exit the atmosphere and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 170 if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere with such a small satellite. * rocket launch speed, which is close to the speed of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|high-altitude flight strategy]], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical component, likely consistent with the rocket flight path initial vector. * final velocity of the rocket, which is the same than the satellite velocity in our SSTO rocket case. At an altitude of 200km, the orbital speed is 7.8 km/s. ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation4.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta v doesn't take into account gravity and other forces acting on the vehicle. * the exhaust velocity of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the best most efficient rocket engine, * thrust is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta v, but it is related to the mass of the payload that has to be accelerated. Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the dry mass of the rocket, once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket, the more efficient it will be in term of delta V. To make a light rocket, high quality materials, thus expensive, must be used. What takes the most volume in a rocket is the fuel and oxidizer tanks. If tanks are solid enough, and this may be a requirements depending on what they contain, they can be used as part of the structure of the rocket. If they are not solid, a external skeleton has to be build that will undergo all mechanical forces that apply to the vehicle. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. * the mass of the engine itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The trust to mass ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. bafb1c4b1719f5cb57aecf0d330d37edd7c48d6d 324 323 2012-08-13T03:29:45Z Vincent 1 /* Delta V achievement */ image name fix wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important points about rockets launched from balloon or from aircraft, mainly what minimum mass we can expect. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require depends on several factors: * rocket launch altitude, the higher it is, the less delta V is required to face atmospheric drag and to exit the atmosphere and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 170 if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere with such a small satellite. * rocket launch speed, which is close to the speed of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|high-altitude flight strategy]], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical component, likely consistent with the rocket flight path initial vector. * final velocity of the rocket, which is the same than the satellite velocity in our SSTO rocket case. At an altitude of 200km, the orbital speed is 7.8 km/s. ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta v doesn't take into account gravity and other forces acting on the vehicle. * the exhaust velocity of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the best most efficient rocket engine, * thrust is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta v, but it is related to the mass of the payload that has to be accelerated. Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the dry mass of the rocket, once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket, the more efficient it will be in term of delta V. To make a light rocket, high quality materials, thus expensive, must be used. What takes the most volume in a rocket is the fuel and oxidizer tanks. If tanks are solid enough, and this may be a requirements depending on what they contain, they can be used as part of the structure of the rocket. If they are not solid, a external skeleton has to be build that will undergo all mechanical forces that apply to the vehicle. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. * the mass of the engine itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The trust to mass ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. 88e911e1bfc54e92b721b1e2aef5873bdd422c6a 327 324 2012-08-14T17:51:55Z Vincent 1 /* Delta V achievement */ example mass ratios wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important points about rockets launched from balloon or from aircraft, mainly what minimum mass we can expect. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require depends on several factors: * rocket launch altitude, the higher it is, the less delta V is required to face atmospheric drag and to exit the atmosphere and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 170 if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere with such a small satellite. * rocket launch speed, which is close to the speed of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|high-altitude flight strategy]], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical component, likely consistent with the rocket flight path initial vector. * final velocity of the rocket, which is the same than the satellite velocity in our SSTO rocket case. At an altitude of 200km, the orbital speed is 7.8 km/s. ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta v doesn't take into account gravity and other forces acting on the vehicle. * the exhaust velocity of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the best most efficient rocket engine, * thrust is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta v, but it is related to the mass of the payload that has to be accelerated. Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the dry mass of the rocket, once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket, the more efficient it will be in term of delta V. To make a light rocket, high quality materials, thus expensive, must be used. What takes the most volume in a rocket is the fuel and oxidizer tanks. If tanks are solid enough, and this may be a requirements depending on what they contain, they can be used as part of the structure of the rocket. If they are not solid, a external skeleton has to be build that will undergo all mechanical forces that apply to the vehicle. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. * the mass of the engine itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The trust to mass ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which may be close to impossible. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still very hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' 99bf5542be39039ae891925d8a5a0d4779a61695 328 327 2012-08-15T03:40:12Z Vincent 1 /* Delta V requirements */ fixes and adding forgotten latitude wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important points about rockets launched from balloon or from aircraft, mainly what minimum mass we can expect. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere with such a small satellite. * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. * '''final velocity of the rocket''', which is the same than the satellite velocity, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta v doesn't take into account gravity and other forces acting on the vehicle. * the exhaust velocity of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the best most efficient rocket engine, * thrust is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta v, but it is related to the mass of the payload that has to be accelerated. Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the dry mass of the rocket, once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket, the more efficient it will be in term of delta V. To make a light rocket, high quality materials, thus expensive, must be used. What takes the most volume in a rocket is the fuel and oxidizer tanks. If tanks are solid enough, and this may be a requirements depending on what they contain, they can be used as part of the structure of the rocket. If they are not solid, a external skeleton has to be build that will undergo all mechanical forces that apply to the vehicle. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. * the mass of the engine itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The trust to mass ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which may be close to impossible. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still very hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' a13ebf23f68b41d479202665f67ce0dfd02a6667 329 328 2012-08-16T01:22:42Z Vincent 1 /* Delta V achievement */ finishing a forgotten sentence ending. wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important points about rockets launched from balloon or from aircraft, mainly what minimum mass we can expect. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere with such a small satellite. * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. * '''final velocity of the rocket''', which is the same than the satellite velocity, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta v doesn't take into account gravity and other forces acting on the vehicle. * the exhaust velocity of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle. * thrust is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta v, but it is related to the mass of the payload that has to be accelerated. Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the dry mass of the rocket, once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket, the more efficient it will be in term of delta V. To make a light rocket, high quality materials, thus expensive, must be used. What takes the most volume in a rocket is the fuel and oxidizer tanks. If tanks are solid enough, and this may be a requirements depending on what they contain, they can be used as part of the structure of the rocket. If they are not solid, a external skeleton has to be build that will undergo all mechanical forces that apply to the vehicle. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. * the mass of the engine itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The trust to mass ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which may be close to impossible. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still very hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' 5624a8f16e85370528063ada3c87b621f9f0722c 331 329 2012-08-16T02:05:05Z Vincent 1 thrust to weight ratio wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important points about rockets launched from balloon or from aircraft, mainly what minimum mass we can expect. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere with such a small satellite. * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. * '''final velocity of the rocket''', which is the same than the satellite velocity, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta v doesn't take into account gravity and other forces acting on the vehicle. * the exhaust velocity of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle. * thrust is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta v, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. In an air-to-orbit launch, the thrust to weight ratio can be lower than for a ground launch, since the initial trajectory angle is not 90°, [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html as explained here] (point 10). They state that a 1.25 thrust to weight ratio is good. * the dry mass of the rocket, once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket, the more efficient it will be in term of delta V. To make a light rocket, high quality materials, thus expensive, must be used. What takes the most volume in a rocket is the fuel and oxidizer tanks. If tanks are solid enough, and this may be a requirements depending on what they contain, they can be used as part of the structure of the rocket. If they are not solid, a external skeleton has to be build that will undergo all mechanical forces that apply to the vehicle. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. * the mass of the engine itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The trust to mass ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which may be close to impossible. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still very hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ce5fe3734bfa3e5cf9d992d59db1ec4dae40924a 332 331 2012-08-16T02:20:04Z Vincent 1 /* Delta V achievement */ bold keywords, few fixes wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important points about rockets launched from balloon or from aircraft, mainly what minimum mass we can expect. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere with such a small satellite. * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. * '''final velocity of the rocket''', which is the same than the satellite velocity, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta v doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta v, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. In an air-to-orbit launch, the thrust to weight ratio can be lower than for a ground launch, since the initial trajectory angle is not 90°, as explained [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html here (point 10)]. They state that a 1.25 thrust to weight ratio is good. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket, the more efficient it will be in term of delta V. To make a light rocket, high quality materials, thus expensive, must be used. What takes the most volume in a rocket is the fuel and oxidizer tanks. If tanks are solid enough, and this may be a requirements depending on what they contain, they can be used as part of the structure of the rocket. If they are not solid, a external skeleton has to be build that will undergo all mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is importantly reduced when launching from altitude than from ground, so the structure can be slightly lighter. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust to weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' c72d5fc9c12f6045745a6e0640c7a082c8b4bd69 333 332 2012-08-18T03:20:52Z Vincent 1 /* Delta V achievement */ orca rocket links wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important points about rockets launched from balloon or from aircraft, mainly what minimum mass we can expect. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere with such a small satellite. * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. * '''final velocity of the rocket''', which is the same than the satellite velocity, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta v doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta v, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. In an air-to-orbit launch, the thrust to weight ratio can be lower than for a ground launch, since the initial trajectory angle is not 90°, as explained [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html here (point 10)]. They state that a 1.25 thrust to weight ratio is good. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket, the more efficient it will be in term of delta V. To make a light rocket, high quality materials, thus expensive, must be used. What takes the most volume in a rocket is the fuel and oxidizer tanks. If tanks are solid enough, and this may be a requirements depending on what they contain, they can be used as part of the structure of the rocket. If they are not solid, a external skeleton has to be build that will undergo all mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is importantly reduced when launching from altitude than from ground, so the structure can be slightly lighter. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust to weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. However, the engine supports only 3 minutes of operation, so it won't probably be enough to reach orbital speed. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' 640c264a32fda092858626746e6a2e1cfae0be71 334 333 2012-08-18T03:49:31Z Vincent 1 /* Delta V requirements */ gravity drag estimation link and first delta V sums wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important points about rockets launched from balloon or from aircraft, mainly what minimum mass we can expect. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles. It still has to be calculated, but if we try to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. * '''final velocity of the rocket''', which is the same than the satellite velocity, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 50 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 200 (earth rotation) = 8080m/s * For choice 2: 7800 (orbital speed) + 70 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 200 (earth rotation) = 7820m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta v doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta v, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. In an air-to-orbit launch, the thrust to weight ratio can be lower than for a ground launch, since the initial trajectory angle is not 90°, as explained [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html here (point 10)]. They state that a 1.25 thrust to weight ratio is good. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket, the more efficient it will be in term of delta V. To make a light rocket, high quality materials, thus expensive, must be used. What takes the most volume in a rocket is the fuel and oxidizer tanks. If tanks are solid enough, and this may be a requirements depending on what they contain, they can be used as part of the structure of the rocket. If they are not solid, a external skeleton has to be build that will undergo all mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is importantly reduced when launching from altitude than from ground, so the structure can be slightly lighter. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust to weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. However, the engine supports only 3 minutes of operation, so it won't probably be enough to reach orbital speed. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' 598153f3b71be321a7ebdcdd23b3d589b6eb5fbb 335 334 2012-08-27T02:00:34Z Vincent 1 ref to Francis's master and adding new minimal mass section wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref>Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. * '''final velocity of the rocket''', which is the same than the satellite velocity, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 50 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 200 (earth rotation) = 8080m/s * For choice 2: 7800 (orbital speed) + 70 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 200 (earth rotation) = 7820m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta v doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta v, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. In an air-to-orbit launch, the thrust to weight ratio can be lower than for a ground launch, since the initial trajectory angle is not 90°, as explained [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html here (point 10)]. They state that a 1.25 thrust to weight ratio is good. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket, the more efficient it will be in term of delta V. To make a light rocket, high quality materials, thus expensive, must be used. What takes the most volume in a rocket is the fuel and oxidizer tanks. If tanks are solid enough, and this may be a requirements depending on what they contain, they can be used as part of the structure of the rocket. If they are not solid, a external skeleton has to be build that will undergo all mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is importantly reduced when launching from altitude than from ground, so the structure can be slightly lighter. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust to weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. However, the engine supports only 3 minutes of operation, so it won't probably be enough to reach orbital speed. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. The reference 1 at bottom addresses this subject. ===Rocket engine=== Since we will design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. We just have to fix a thrust-to-weight ratio for the engine, so that we can fix it when the rest of the vehicle's mass is evaluated. The thrust-to-weight ratio for the entire rocket will also be required, as it will determine what will be the initial acceleration. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature, a ratio of 2.0 is common for small vehicles. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here, as explained in paper 1 too. A ratio of more than 100 is common in small launchers' engines. ===Fuel tanks=== Fuel tanks mass depend mostly on the type of fuel delivery system used: pressurization or pumping. When using pressurized tanks to force the fuel into the engine, the tanks have to sustain the pressure and must have thicker walls. A pumped fuel only requires the fuel to not leak into vacuum, since we are going to space. For kerosene-like fuel, it can be very simple. For the cryogenic oxidizer, it's more complicated due to the evaporation losses related to poor isolation and to metal weakening with large temperature differential. ===Structure=== Fuel tanks will be the largest part of the vehicle, using them as structure may save mass if thought carefully. ===Avionics=== ===Other elements=== GIMBAL, Valves, pressurization system, RCS, communication, fairing, separation system at least. ==References== <references /> 8a73a210f5ba5891c15f2dd0d9b483e490c38f19 336 335 2012-08-27T02:59:23Z Vincent 1 fixing multiple references and moving text in related new subsections wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. The reference <ref="francis" /> addresses this subject. ===Rocket engine=== Since we will design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. We just have to fix a thrust-to-weight ratio for the engine, so that we can fix it when the rest of the vehicle's mass is evaluated. The thrust-to-weight ratio for the entire rocket will also be required, as it will determine what will be the initial acceleration. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. ===Fuel tanks=== Fuel tanks mass depend mostly on the type of fuel delivery system used: pressurization or pumping. When using pressurized tanks to force the fuel into the engine, the tanks have to sustain the pressure and must have thicker walls. A pumped fuel only requires the fuel to not leak into vacuum, since we are going to space. For kerosene-like fuel, it can be very simple. For the cryogenic oxidizer, it's more complicated due to the evaporation losses related to poor isolation and to metal weakening with large temperature differential. ===Structure=== What takes the most volume in a rocket is the fuel and oxidizer tanks. If tanks are solid enough, they can be used as part of the structure of the rocket. If they are not solid, a external skeleton has to be build that will undergo all mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. ===Avionics=== <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. ===Other elements=== GIMBAL, Valves, pressurization system, RCS, communication, fairing, separation system at least. ==References== <references /> 5a38c74f01b20fd162acaafc6ce9ee85cf6d8bd0 337 336 2012-09-03T17:46:39Z Vincent 1 /* Systems design for minimum mass */ fixing reference wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. The reference <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. We just have to fix a thrust-to-weight ratio for the engine, so that we can fix it when the rest of the vehicle's mass is evaluated. The thrust-to-weight ratio for the entire rocket will also be required, as it will determine what will be the initial acceleration. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. ===Fuel tanks=== Fuel tanks mass depend mostly on the type of fuel delivery system used: pressurization or pumping. When using pressurized tanks to force the fuel into the engine, the tanks have to sustain the pressure and must have thicker walls. A pumped fuel only requires the fuel to not leak into vacuum, since we are going to space. For kerosene-like fuel, it can be very simple. For the cryogenic oxidizer, it's more complicated due to the evaporation losses related to poor isolation and to metal weakening with large temperature differential. ===Structure=== What takes the most volume in a rocket is the fuel and oxidizer tanks. If tanks are solid enough, they can be used as part of the structure of the rocket. If they are not solid, a external skeleton has to be build that will undergo all mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. ===Avionics=== <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. ===Other elements=== GIMBAL, Valves, pressurization system, RCS, communication, fairing, separation system at least. ==References== <references /> ad71393bcc2ccbe02eb3a694d83b2ecea003005f 343 337 2012-10-28T00:19:12Z Vincent 1 /* Systems design for minimum mass */ more about mass for all subsystems wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Since a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, an even more powerful engine is required and this is an iterative process. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the wall's thickness and material are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the vehicle will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. ===Fuel tanks=== Fuel tanks mass depend mostly on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine, the tanks have to sustain the pressure and must have thicker walls. When pumping fuel, the tank has to prevent it to leak into vacuum. For kerosene-like fuel, it can be very simple. For the cryogenic oxidizer, it's more complicated due to the evaporation losses related to poor isolation, thus requiring pressurization anyway, and to metal weakening with large temperature differential. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the dynamic pressure stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. '''TODO''': Weight_tank = f_Al(volume, internal pressure) ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. ===Other elements=== GIMBAL, Valves, pressurization system, RCS, communication, fairing, orbit insertion mechanism, and so on. ==References== <references /> 7a21345fa82c851e1570ff6523d98ae5e912c747 345 343 2012-10-28T04:22:06Z Vincent 1 /* Systems design for minimum mass */ more mass-related text wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Since a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, an even more powerful engine is required and this is an iterative process. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the wall's thickness and material are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. ===Fuel tanks=== Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine, the tanks have to sustain the pressure and must have thicker walls. When pumping fuel, the tank has to prevent it to leak into vacuum. For kerosene-like fuel, it can be very simple. For the [[LOX|cryogenic oxidizer]], it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, thus requiring pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the dynamic pressure stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. '''TODO''': Weight_tank = f_Al(volume, internal pressure) ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==References== <references /> 20fae20a01803700302a543c8a83f8860c0d3415 346 345 2012-11-01T01:19:55Z Vincent 1 /* Delta V achievement */ fix formulas images color wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Since a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, an even more powerful engine is required and this is an iterative process. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the wall's thickness and material are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. ===Fuel tanks=== Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine, the tanks have to sustain the pressure and must have thicker walls. When pumping fuel, the tank has to prevent it to leak into vacuum. For kerosene-like fuel, it can be very simple. For the [[LOX|cryogenic oxidizer]], it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, thus requiring pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the dynamic pressure stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. '''TODO''': Weight_tank = f_Al(volume, internal pressure) ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==References== <references /> 4a9bdbb9e020ed7fcff8645e6f0b1caf52b956d3 0 1 325 294 2012-08-14T16:42:57Z Vincent 1 /* The rocket */ adding links to rocket pages wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, you can [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide a low cost orbital launch system for small size satellites, probably with a mass lower than 1kg. Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. The rocket, as in any other orbital launch system, would make it to orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. ==News== '''''May 21, 2012: ''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may reflect some weak or false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design it the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. Can electricity energy be considered for that kind of mission? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, a major issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive and their use is restricted to missiles or UAVs. ===Staging and recovery=== Separation from the rocket is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a complicated autonomous return-to-runway and landing system. ===Guidance=== A satellite navigation system can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. I believe that mere cameras can be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. Anyway, if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: 1) the rocket engine has to be directionally controllable (generally using hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]), or 2) control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) must be used to control the attitude of the rocket, as partially does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but are mandatory in our case. This is one of the big differences between sub-orbital and orbital space flight. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 98621ac05952f8c394f48a3a2bd3849fddf684eb 326 325 2012-08-14T17:00:49Z Vincent 1 /* News */ august: work on blade manufacturing process wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, you can [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide a low cost orbital launch system for small size satellites, probably with a mass lower than 1kg. Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. The rocket, as in any other orbital launch system, would make it to orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. ==News== '''''August 2012 update: ''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012: ''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may be weakly verified or partly false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design it the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. Can electricity energy be considered for that kind of mission? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, a major issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive and their use is restricted to missiles or UAVs. ===Staging and recovery=== Separation from the rocket is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a complicated autonomous return-to-runway and landing system. ===Guidance=== A satellite navigation system can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. I believe that mere cameras can be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. Anyway, if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: 1) the rocket engine has to be directionally controllable (generally using hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]), or 2) control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) must be used to control the attitude of the rocket, as partially does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but are mandatory in our case. This is one of the big differences between sub-orbital and orbital space flight. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. e2e9be331359ca7e1876ae6747cb15fead5b841a 0 11 338 265 2012-10-15T02:20:55Z Vincent 1 https for wikipedia wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets or other engines. They are however very difficult to design and manufacture and are thus very expensive. On this page, we will explore how costs can be reduced while still having adequate capabilities for high altitude flight. ==General principles== Lots of information are available on [https://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that puts energy into a gas, this energy is extracted by a turbine, and the turbine drives both the fan that provides thrust and the compression stage that feeds the combustion with oxygen. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake, and so on. The fan provides thrust by creating a massive air flow, and the engine's core also creates thrust by evacuating the high-speed hot combustion gas. In commercial turbofan engines, the fan is generally responsible for 90% of the overall thrust. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [https://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost, has the fan mounted on it. One stage engines exist and are less complicated and expensive to build, but are much less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan, but this is not an easy task due to this very speed. Commercial engines featuring a gearbox for the turbofan's fan are expected to reach market in 2012. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan, but are not yet optimal. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seeing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down with the cost. ===Shaped core or shaped shaft?=== An obvious but important optimization to reduce cost and complexity of manufacturing is to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to design and produce, with a simple [https://en.wikipedia.org/wiki/Lathe lathe] ([https://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [https://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates, for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a shaft turned in raw metal won't change much on its final mass. Moreover, we may use a turbine-level mechanism embedded in the stator to try to cool it, which would make it hollow. The main mechanical issues are probably how to properly fix the blades on rotor and stator, how to fix the rotor on the stator with little gap, and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compressor blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [https://en.wikipedia.org/wiki/Superalloys superalloys] or are ceramic-coated. It's the inability of blades to withstand heat and work that limit the power of the engine. Indeed, around 70% of the gas provided by the compressor is used only for chamber and turbine cooling, instead of using it to burn more fuel and create more thrust. The compressor and the turbine are not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow driven by the action of rotor blades to form inside the engine, which would decrease the energy of the gas. Stator blades or vanes redirect the airflow on the next stage in the more efficient direction. Highest efficiencies are reached in turbofans when gaps are reduced between rotor blades' tip and the stator, as well as between the stator blades' tip and the rotor. As always, good efficiency means high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. The shape of the blade and the parameters of their cascade also affects the efficiency. A small 5 stage supersonic compressor providing the same pressure rise than a 15 stage subsonic compressor is less efficient, but it may be compensated by the higher thrust-to-weight ratio. Blade geometric design is also very complicated. First turbine engines had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1926, [https://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines Alan A. Griffith] proved that if blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to deserve being built. Airfoils for blade designs allow compressor stages to better increase the static pressure since they create an expander, an increasing area for the air flow to pass through. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C so it will be way off-limits, and even below that, it will deform before melting. And should we mention that blade deformation at high-centrifugal loads, caused by the high rpm, is a very good way to blow the engine off? Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the vanes the turbine, which are, on actual engines, made of a ceramic-coated high-temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge or bleed circulates, enters the blades, and evacuates through small holes in the blades (convective cooling and film cooling). For the rotor blades, the same principle is used, but with compressor air passing inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air or a high speed electric engine for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air and compressor stall, re-ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude. The rotor speed information and altimeter may be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and removed axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we aim. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, a simple square-section root is probably enough. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blade-roots are inserted and fixed. Creating a perpendicular blade root is already a challenge. Rotor blades would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting should be considered. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Air bled from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose [[Turbofan:Alternative Designs|alternative turbofan designs]] based on a mixed-flow turbine. Several pages have been created in the [[:Category:Turbofan|Turbofan category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three to five stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while sustaining the combustion. * [[Turbofan:Bearings|Bearings and cooling]]: high speed rotations require adapted bearings and cooling, which may be reused for rotor and even turbine cooling. ===Turbofan design procedure=== # evaluate required thrust (from aircraft mass and lift, but also [[Flight_at_high_altitude|flight characteristics]]) # calculate required mass flow rate for the fan (thust is [https://en.wikipedia.org/wiki/Thrust calculated] from MFR and flow speed) # fix bypass ratio and fan diameter and rpm, thus giving core diameter (BPR may be [[Turbofan:Alternative_Designs#Full_transonic_engine_design_in_a_single_spool_with_2.1_BPR|fixed by design]]) # calculate required power to drive the fan alone # evaluate a gross compressor driving power (refined later)<br />&nbsp; # calculate total power that has to be drawn from the turbine (fan + compressor + losses) # calculate mass flow rate for the combustion alone # calculate mass flow rate for cooling chamber and turbine ## calculate mass flow rate for cooling chamber ## evaluate mass flow rate for cooling turbine to add to the latter ## calculate entropy and fluid parameters at combustor discharge (speed, temperature) ## calculate temperature of turbine vanes and blade and check if it is acceptable ## iterate on item 8.1 until temperature is unacceptable # calculate the number of turbine blades and stages required for this power # calculate compressor discharge pressure and pressure ratio # calculate how many compressor stages are required depending on sonic or supersonic blade design and fix design # calculate compressor driving power # iterate on item 6 until total power varies # design blades for all calculated parameters and re-run at item 6, total power may have changed [[Category:Turbofan]] 89b9a7b3e31de3cc85a58bbdfebb7953c3dd733e 0 47 339 230 2012-10-15T03:05:52Z Vincent 1 text fixes and https wikipedia links wikitext text/x-wiki =Bearings and cooling= Rotational speed achieved by the engine will probably be above 40000rpm. At these speeds, regular ball bearings may overheat or suffer from a too fast wear. In real turbine engines, bearings are constantly lubricated by oil jets, which poses problems with regards to oil pressurization and leaks in other parts of the engine. Seals are consequently placed close to bearings to prevent leaks, generally carbon seals. To enforce the seal, oil is contained in a casing with an internal pressure lower than the external pressure built from compressed air. That way, air can enter the oil casing, but oil cannot leak outside, in other parts of the engine. Accessories like oil pumps, pipes, fixations, filters, tanks, heat exchangers and so on, are also required. ==Bearings== ===Ball bearings=== Ball or roller bearings are the obvious way to guide rotating parts. They can handle high mechanical constraints radially or even axially, they are inexpensive and their integration is reasonably simple. [https://en.wikipedia.org/wiki/Silicon_nitride#Bearings Silicon nitride bearings] have lots of improvements over regular metal ball bearings. Balls are more than 60% less heavy, thus having a lower inertia at high speeds, implying a more softer contact with the tracks, allowing longer lifetime or higher reachable speeds. They also require less lubrication. Fortunately, silicon nitride bearings have reached market with a large production, and are not over-expensive. ===Alternate bearings=== Fluid or magnetic bearings should be considered. They allow much higher rotation speeds and lower friction, but have two main drawbacks. At standby state, they release the radial constrain on moving parts. In reduced-size turbomachinery, where rotor and stator have to be adjusted to tens of microns, it seems quite complicated to use those bearings. The second drawback is that they require more external hardware, to pressurize the fluid or to provide magnetic energy. However, magnetic bearings have been demonstrated in [http://books.google.com/books?id=AXtqMugS3TQC&lpg=PP1&pg=PA263#v=onepage&q&f=false this paper] <ref>S. Jana, V. Arun Kumar and M. Ananda. '''5-axes levitation of a rotor towards indigenization of the magnetic bearing technology'''. In ''Air breathing engines and aerospace propulsion: proceedings of NCABE 2004'', november 2004.</ref>, in which axial position accuracy is measured below 150µm for a 4kg rotor at around 2000rpm. The rotor position sensor has a resolution of 2µm per mV. Unfortunately, no indication is given about the resting position of the rotor and how it impacts the clearance between rotor and stator. [https://en.wikipedia.org/wiki/Foil_bearing Foil bearings] are a particular type of fluid bearing, that "Unlike aero or hydrostatic bearings, foil bearings require no external pressurisation system for the working fluid, so the hydrodynamic bearing is self-starting". In [http://b-dig.iie.org.mx/BibDig/P06-0351/pdfs/track-16/GT2006-90791.pdf this other paper] <ref>Hooshang Heshmat, Michael J. Tomaszewski, James F. Walton II. '''Small gas turbine engine operating with high temperature foil bearing'''. In ''proceedings of GT2006 ASME Turbo Expo 2006: Power for land, sea and air'', may 2006.</ref>, a small centrifugal turbojet is built to evaluate the ability of [http://www.miti.cc/products-services.html MiTi]'s product, a foil bearing, to sustain very high rotation speeds (120'000rpm) and high temperature (800°C). The bearing has a low spacing between the rotor's journal and the stator fixation, but it is secured, in this paper, using a ball bearing on the compressor side, where the temperature is low. They planned to make a dual-foil bearing, we'll need to check on that. MiTi also demonstrated a [http://www.miti.cc/newsletters/20_150mm_foil_journal_bearing%20_hybrid_foil_magnetic_bearing.pdf hybrid foil magnetic bearing], that has the advantages of magnetic bearings at low speeds and those of foil bearings at high speeds. ==Use of lubricating oil for cooling== In real-world jet engines, cooling is the primary function of oil when conventional bearings are used, even more important than lubrication. That's well explained in [https://www.youtube.com/watch?v=WAia8PwMvQM this AgentJayz video]. A high flow rate of oil is then required, with a heat exchanger somewhere along the oil path. The other fluid for the heat exchanger can be air from the bypass duct or fuel, but in our highly size-constrained engine environment, we'll probably have to move some of the engine's equipment to the wings. But that will have to be studied after the bearing type has been chosen obviously. ==Oil displacement without external pumping== [https://en.wikipedia.org/wiki/Screw_conveyor Screw pumping] may be used as a way to move the oil through the engine, i.e. parts that have to be lubricated and cooled. The idea is to use the work of the turbine to directly drive the oil pumping without requiring external accessories. That does not solve the fact that cooling the oil may require external hardware, and that sealing is mandatory. ==Other hardware required for lubrication and bearing cooling== Sensors will be required too, at least for oil temperature and displacement confirmation. Oil temperature informs about the status of the engine's bearings. Oil displacement sensor is required to ensure that there is no problem with the oil/cooling flow in the engine and that the measured temperature is not bogus. Simple oil filters should also be put somewhere on the oil lines to prevent the more obvious failures. ==References== <references /> [[Category:Turbofan|Bearings and cooling]] 144cfc063ec3c62f4ff8ecd2ba79a2ca1bf68848 341 339 2012-10-23T01:42:59Z Vincent 1 /* Oil displacement without external pumping */ discussion about the method wikitext text/x-wiki =Bearings and cooling= Rotational speed achieved by the engine will probably be above 40000rpm. At these speeds, regular ball bearings may overheat or suffer from a too fast wear. In real turbine engines, bearings are constantly lubricated by oil jets, which poses problems with regards to oil pressurization and leaks in other parts of the engine. Seals are consequently placed close to bearings to prevent leaks, generally carbon seals. To enforce the seal, oil is contained in a casing with an internal pressure lower than the external pressure built from compressed air. That way, air can enter the oil casing, but oil cannot leak outside, in other parts of the engine. Accessories like oil pumps, pipes, fixations, filters, tanks, heat exchangers and so on, are also required. ==Bearings== ===Ball bearings=== Ball or roller bearings are the obvious way to guide rotating parts. They can handle high mechanical constraints radially or even axially, they are inexpensive and their integration is reasonably simple. [https://en.wikipedia.org/wiki/Silicon_nitride#Bearings Silicon nitride bearings] have lots of improvements over regular metal ball bearings. Balls are more than 60% less heavy, thus having a lower inertia at high speeds, implying a more softer contact with the tracks, allowing longer lifetime or higher reachable speeds. They also require less lubrication. Fortunately, silicon nitride bearings have reached market with a large production, and are not over-expensive. ===Alternate bearings=== Fluid or magnetic bearings should be considered. They allow much higher rotation speeds and lower friction, but have two main drawbacks. At standby state, they release the radial constrain on moving parts. In reduced-size turbomachinery, where rotor and stator have to be adjusted to tens of microns, it seems quite complicated to use those bearings. The second drawback is that they require more external hardware, to pressurize the fluid or to provide magnetic energy. However, magnetic bearings have been demonstrated in [http://books.google.com/books?id=AXtqMugS3TQC&lpg=PP1&pg=PA263#v=onepage&q&f=false this paper] <ref>S. Jana, V. Arun Kumar and M. Ananda. '''5-axes levitation of a rotor towards indigenization of the magnetic bearing technology'''. In ''Air breathing engines and aerospace propulsion: proceedings of NCABE 2004'', november 2004.</ref>, in which axial position accuracy is measured below 150µm for a 4kg rotor at around 2000rpm. The rotor position sensor has a resolution of 2µm per mV. Unfortunately, no indication is given about the resting position of the rotor and how it impacts the clearance between rotor and stator. [https://en.wikipedia.org/wiki/Foil_bearing Foil bearings] are a particular type of fluid bearing, that "Unlike aero or hydrostatic bearings, foil bearings require no external pressurisation system for the working fluid, so the hydrodynamic bearing is self-starting". In [http://b-dig.iie.org.mx/BibDig/P06-0351/pdfs/track-16/GT2006-90791.pdf this other paper] <ref>Hooshang Heshmat, Michael J. Tomaszewski, James F. Walton II. '''Small gas turbine engine operating with high temperature foil bearing'''. In ''proceedings of GT2006 ASME Turbo Expo 2006: Power for land, sea and air'', may 2006.</ref>, a small centrifugal turbojet is built to evaluate the ability of [http://www.miti.cc/products-services.html MiTi]'s product, a foil bearing, to sustain very high rotation speeds (120'000rpm) and high temperature (800°C). The bearing has a low spacing between the rotor's journal and the stator fixation, but it is secured, in this paper, using a ball bearing on the compressor side, where the temperature is low. They planned to make a dual-foil bearing, we'll need to check on that. MiTi also demonstrated a [http://www.miti.cc/newsletters/20_150mm_foil_journal_bearing%20_hybrid_foil_magnetic_bearing.pdf hybrid foil magnetic bearing], that has the advantages of magnetic bearings at low speeds and those of foil bearings at high speeds. ==Use of lubricating oil for cooling== In real-world jet engines, cooling is the primary function of oil when conventional bearings are used, even more important than lubrication. That's well explained in [https://www.youtube.com/watch?v=WAia8PwMvQM this AgentJayz video]. A high flow rate of oil is then required, with a heat exchanger somewhere along the oil path. The other fluid for the heat exchanger can be air from the bypass duct or fuel, but in our highly size-constrained engine environment, we'll probably have to move some of the engine's equipment to the wings. But that will have to be studied after the bearing type has been chosen obviously. ==Oil displacement without external pumping== [https://en.wikipedia.org/wiki/Screw_conveyor Screw pumping] is being assessed as the way to displace the oil throughout the engine. There will probably be two bearings in the engine, both requiring an oil bath and the shaft itself can probably also use a little refreshment. The idea is to use the rotation of the shaft to actually displace the oil without requiring external accessories. That would be a very lightweight solution and perhaps not that hard to implement, since our shaft is not hollow. The principle first has to be verified, then be tested with such high rotation rates in order to verify that the drag generated on the shaft is acceptable. That does not solve the fact that cooling the oil requires external hardware (a heat exchanger), and that sealing is mandatory in the oil inlet and outlet areas, generally where the bearings are. ==Other hardware required for lubrication and bearing cooling== Sensors will be required too, at least for oil temperature and displacement confirmation. Oil temperature informs about the status of the engine's bearings. Oil displacement sensor is required to ensure that there is no problem with the oil/cooling flow in the engine and that the measured temperature is not bogus. Simple oil filters should also be put somewhere on the oil lines to prevent the more obvious failures. ==References== <references /> [[Category:Turbofan|Bearings and cooling]] f1c36ffd3aa1e75eb4edccb9adffe18d08931b6d 0 17 342 122 2012-10-27T20:36:20Z Vincent 1 /* Cooling for a LOX/E85 engine */ typo wikitext text/x-wiki =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, instead of fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine failure. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotary Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Moreover, LOX and cheap fuels are readily available. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. From SpaceX's experience, it seems that regenerative cooling is sufficient to sustain the combustion temperature hitting the walls. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. fe8a9f941187fdcf5293d33da8ed6732806cbc88 348 342 2012-11-01T01:29:39Z Vincent 1 /* Rocket Engine */ page summary and link to the approximations wikitext text/x-wiki This page is an introduction to rocket engines, explaining and comparing some of its subsystems. Our rocket will be based on regular rocket technology, as explained on this page. However some choices need to be made in order to gauge the feasibility of the project as a whole, in particular to have ideas of the possible dimensions of the rocket. These choices are presented on the page of the [[Rocket:First approximations|first approximations for the rocket]]. =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, instead of fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine failure. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotary Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Moreover, LOX and cheap fuels are readily available. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. From SpaceX's experience, it seems that regenerative cooling is sufficient to sustain the combustion temperature hitting the walls. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. 9bfd727da280e9a8a3f54372d2da7bde5e5430b6 0 62 344 2012-10-28T03:33:29Z Vincent 1 about LOX wikitext text/x-wiki =Liquid Oxygen= See [https://en.wikipedia.org/wiki/Oxygen Wikipedia's page]. Boiling point of -182.95°C (90.20 K). Critical temperature of -118.59°C (154.59 K) meaning that it cannot stay liquid above this temperature. 1 Litre of LOX provides 840 Litre of GOX (gaseous oxygen) at 1atm/20°C. Like any liquid, LOX vaporizes when stored above its boiling point. The vapour builds up to pressure that the tank will no be able to contain, so a relief valve must be put in place to compensate this evaporation. Insulation limits the rate of vaporization, the best being the vacuum space between a tank inner and outer walls, like Dewar flasks. fa4363ddf922c5f804699016ee17fb565df58957 6 63 349 2012-11-01T04:43:33Z Vincent 1 Graph representing rocket dry and wet mass as function of the payload mass (or any other fixed mass subsystem). wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the payload mass (or any other fixed mass subsystem). 92472815e8c754715a7c221ae990a448ba69138f 6 64 350 2012-11-01T04:44:48Z Vincent 1 Graph representing rocket dry and wet mass as function of the mass ratio (fixed by Isp and Delta V). wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the mass ratio (fixed by Isp and Delta V). c170bd9fa96930e3f5461fb37c874b24caedefe4 6 65 351 2012-11-01T04:45:32Z Vincent 1 Graph representing rocket dry and wet mass as function of the engine's thrust-to-weight ratio. wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the engine's thrust-to-weight ratio. 629900940158a61aadb92c7445d9fa61534134dc 6 66 352 2012-11-01T04:47:26Z Vincent 1 Graph representing rocket dry and wet mass as function of the vehicle (overall) initial thrust-to-weight ratio. There is an equivalence with the acceleration in number of G's. wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the vehicle (overall) initial thrust-to-weight ratio. There is an equivalence with the acceleration in number of G's. b2d2d826fab4dc95c2a9f449df1c3c2b38c484f8 0 61 353 346 2012-11-01T05:12:09Z Vincent 1 algorithm for the iterative process and first graphs wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Since a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, an even more powerful engine is required and this is an iterative process. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the wall's thickness and material are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. ===Fuel tanks=== Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine, the tanks have to sustain the pressure and must have thicker walls. When pumping fuel, the tank has to prevent it to leak into vacuum. For kerosene-like fuel, it can be very simple. For the [[LOX|cryogenic oxidizer]], it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, thus requiring pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the dynamic pressure stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. '''TODO''': Weight_tank = f_Al(volume, internal pressure) ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The code is coming soon online. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. The variation of one parameter is represented in the graphs below, emphasizing the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. [[File:Payload_0-80kg.png]] [[File:Mass_ratio.png]] The first graph was made in the first place with larger values for the payload mass, but it looked like it was linear. I zoomed on the lower mass part since it's more our concern here, and to emphasize the slight non-linearity for small satellites. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. The second graph is the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Engine_thrust_to_weight.png]] [[File:Overall_thrust_to_weight.png]] These graphs highlight the importance of the thrust-to-weight ratios. The first is related to an engine design issue, the second is related to the rocket's trajectory. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.2. ==References== <references /> 6a975f2e2f3e31b8178c241bbcd37b3c93bd06df 356 353 2012-11-02T02:37:25Z Vincent 1 /* Minimum mass evaluation */ fifth graph: thrust and final acceleration, new close up graph for payload mass, and first approximation. wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Since a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, an even more powerful engine is required and this is an iterative process. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the wall's thickness and material are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. ===Fuel tanks=== Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine, the tanks have to sustain the pressure and must have thicker walls. When pumping fuel, the tank has to prevent it to leak into vacuum. For kerosene-like fuel, it can be very simple. For the [[LOX|cryogenic oxidizer]], it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, thus requiring pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the dynamic pressure stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. '''TODO''': Weight_tank = f_Al(volume, internal pressure) ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The code is coming soon online. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 15.2. ''This is the value for a 9km/s Delta V and a constant Isp of 340s (propellants are 93.4% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is full aluminium, even skirts and fairing. '''Results:''' * '''DRY MASS: 28.3533 kg''', WET mass: 430.969 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the rocket's diameter): 3.80296 x 0.380296 m * Engine thrust: 5279.38 N, engine mass: 5.38712 kg * Final thrust-to-weight-ratio: 19 G * Actuators mass: 0.879896 kg, wiring mass: 0.266207 kg, structure mass: 5.79303 kg * LOX tank info: ** diameter: 0.380296 m, length: 2.30944 m ** thickness: 0.002 m, volume: 0.244837 m^3, mass: 14.8995 kg * E85 tank info: ** diameter: 0.380296 m, length: 1.49351 m ** thickness: 0.0005 m, volume: 0.158336 m^3, mass: 0.847568 kg '''So here we are. 431kg to carry up to 30km altitude.''' The numbers do not include any margin for error, meaning that if we count on some LOX venting, we need a larger LOX tank, and again much more mass at the end. Some structural parts like skirts and tanks outer skins (if required) could easily be made of composite material. That could reduce the structure mass probably by 20% (see the payload graph below, a decrease in 1kg of fixed mass is approximately balanced by a decrease of 50 kg in wet mass). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> f372475cd954a25fc8deb79ac5ecd6e7a3ed1686 358 356 2012-11-02T02:54:05Z Vincent 1 /* Minimum mass evaluation */ source code link wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Since a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, an even more powerful engine is required and this is an iterative process. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the wall's thickness and material are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. ===Fuel tanks=== Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine, the tanks have to sustain the pressure and must have thicker walls. When pumping fuel, the tank has to prevent it to leak into vacuum. For kerosene-like fuel, it can be very simple. For the [[LOX|cryogenic oxidizer]], it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, thus requiring pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the dynamic pressure stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. '''TODO''': Weight_tank = f_Al(volume, internal pressure) ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code is available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 15.2. ''This is the value for a 9km/s Delta V and a constant Isp of 340s (propellants are 93.4% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is full aluminium, even skirts and fairing. '''Results:''' * '''DRY MASS: 28.3533 kg''', WET mass: 430.969 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the rocket's diameter): 3.80296 x 0.380296 m * Engine thrust: 5279.38 N, engine mass: 5.38712 kg * Final thrust-to-weight-ratio: 19 G * Actuators mass: 0.879896 kg, wiring mass: 0.266207 kg, structure mass: 5.79303 kg * LOX tank info: ** diameter: 0.380296 m, length: 2.30944 m ** thickness: 0.002 m, volume: 0.244837 m^3, mass: 14.8995 kg * E85 tank info: ** diameter: 0.380296 m, length: 1.49351 m ** thickness: 0.0005 m, volume: 0.158336 m^3, mass: 0.847568 kg '''So here we are. 431kg to carry up to 30km altitude.''' The numbers do not include any margin for error, meaning that if we count on some LOX venting, we need a larger LOX tank, and again much more mass at the end. Some structural parts like skirts and tanks outer skins (if required) could easily be made of composite material. That could reduce the structure mass probably by 20% (see the payload graph below, a decrease in 1kg of fixed mass is approximately balanced by a decrease of 50 kg in wet mass). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> b3f6376487f0dd976330eff63db8c0ac57601a52 366 358 2012-11-06T03:52:33Z Vincent 1 /* Our first approximation */ 1kg payload hint wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Since a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, an even more powerful engine is required and this is an iterative process. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the wall's thickness and material are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. ===Fuel tanks=== Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine, the tanks have to sustain the pressure and must have thicker walls. When pumping fuel, the tank has to prevent it to leak into vacuum. For kerosene-like fuel, it can be very simple. For the [[LOX|cryogenic oxidizer]], it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, thus requiring pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the dynamic pressure stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. '''TODO''': Weight_tank = f_Al(volume, internal pressure) ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code is available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 15.2. ''This is the value for a 9km/s Delta V and a constant Isp of 340s (propellants are 93.4% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is full aluminium, even skirts and fairing. '''Results:''' * '''DRY MASS: 28.3533 kg''', WET mass: 430.969 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the rocket's diameter): 3.80296 x 0.380296 m * Engine thrust: 5279.38 N, engine mass: 5.38712 kg * Final thrust-to-weight-ratio: 19 G * Actuators mass: 0.879896 kg, wiring mass: 0.266207 kg, structure mass: 5.79303 kg * LOX tank info: ** diameter: 0.380296 m, length: 2.30944 m ** thickness: 0.002 m, volume: 0.244837 m^3, mass: 14.8995 kg * E85 tank info: ** diameter: 0.380296 m, length: 1.49351 m ** thickness: 0.0005 m, volume: 0.158336 m^3, mass: 0.847568 kg '''So here we are. 431kg to carry up to 30km altitude.''' These numbers do not include any margin for error. If we count on some LOX venting, we need a larger LOX tank, and again much more mass at the end. Some structural parts like skirts and tanks outer skins (if required) could easily be made of composite material. That could reduce the structure mass probably by 20% (see the payload graph below, a decrease in 1kg of fixed mass is approximately balanced by a decrease of 50 kg in wet mass). We could increase payload to 1kg to allow for more missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> 6e3fc5551b9e2ea8a3f84293caf9d249e076654e 371 366 2012-11-07T03:46:31Z Vincent 1 /* Our first approximation */ 8.0 delta V instead of 9.0, lox venting, composite skirts, foam insulation wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Since a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, an even more powerful engine is required and this is an iterative process. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the wall's thickness and material are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. ===Fuel tanks=== Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine, the tanks have to sustain the pressure and must have thicker walls. When pumping fuel, the tank has to prevent it to leak into vacuum. For kerosene-like fuel, it can be very simple. For the [[LOX|cryogenic oxidizer]], it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, thus requiring pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the dynamic pressure stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. '''TODO''': Weight_tank = f_Al(volume, internal pressure) ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code is available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * '''DRY MASS: 12.8552 kg''', WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg '''So here we are. 142kg to carry up to 30km altitude.''' The LOX venting has not been properly calculated based on insulation and real vaporization heat data. We could increase payload to 1kg to allow for more missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> eef733d67ca0865964629bb6bfaa69cf22ab062c 375 371 2012-11-07T18:41:55Z Vincent 1 /* Minimum mass evaluation */ text update and graph parameters wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Since a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, an even more powerful engine is required and this is an iterative process. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the wall's thickness and material are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. ===Fuel tanks=== Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine, the tanks have to sustain the pressure and must have thicker walls. When pumping fuel, the tank has to prevent it to leak into vacuum. For kerosene-like fuel, it can be very simple. For the [[LOX|cryogenic oxidizer]], it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, thus requiring pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the dynamic pressure stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. '''TODO''': Weight_tank = f_Al(volume, internal pressure) ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * '''DRY MASS: 12.8552 kg''', WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg '''So here we are. 142kg to carry up to 30km altitude.''' The LOX venting has not been properly calculated based on insulation and real vaporization heat data. We could increase payload to 1kg to allow for more missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. Rocket parameters are 15.2 mass ratio (9.0 Delta V, 340 constant I_sp), propellants are LOX and E85 at stoichiometric ratio (no venting, no tank insulation), full aluminium structure. Other parameters are the same than above, and the software used was version 1 (see updates here: [[File:Rocket_mass.c]]). [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> 94fbabe06c57f87b7306ff6f9f5a95920e9f3b95 392 375 2012-11-09T15:31:44Z Vincent 1 /* Our first approximation */ precisions about the current model wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Since a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, an even more powerful engine is required and this is an iterative process. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the wall's thickness and material are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. ===Fuel tanks=== Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine, the tanks have to sustain the pressure and must have thicker walls. When pumping fuel, the tank has to prevent it to leak into vacuum. For kerosene-like fuel, it can be very simple. For the [[LOX|cryogenic oxidizer]], it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, thus requiring pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the dynamic pressure stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. '''TODO''': Weight_tank = f_Al(volume, internal pressure) ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * '''DRY MASS: 12.8552 kg''', WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg '''So here we are. 142kg to carry up to 30km altitude.''' The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. Rocket parameters are 15.2 mass ratio (9.0 Delta V, 340 constant I_sp), propellants are LOX and E85 at stoichiometric ratio (no venting, no tank insulation), full aluminium structure. Other parameters are the same than above, and the software used was version 1 (see updates here: [[File:Rocket_mass.c]]). [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> d6951f2a0a840bd548b7153a154cbc5d3904886f 393 392 2012-11-09T15:45:32Z Vincent 1 /* Fuel tanks */ tanks update and link wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Since a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, an even more powerful engine is required and this is an iterative process. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the wall's thickness and material are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * '''DRY MASS: 12.8552 kg''', WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg '''So here we are. 142kg to carry up to 30km altitude.''' The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. Rocket parameters are 15.2 mass ratio (9.0 Delta V, 340 constant I_sp), propellants are LOX and E85 at stoichiometric ratio (no venting, no tank insulation), full aluminium structure. Other parameters are the same than above, and the software used was version 1 (see updates here: [[File:Rocket_mass.c]]). [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> 14ac2ed2d3b47c10ffe02c84a718a9302a5e4263 395 393 2012-11-09T15:59:11Z Vincent 1 /* Minimum mass evaluation */ forgotten telemetry wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the Isp of the engine, as we can see here: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is very important and we should aim and actual Isp approaching 20s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine, the sea level Isp is generally around 10% less, and since we will use a cheap design, we may not have the optimal nozzle and fuel combustion efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the fuel and consequently Isp have been fixed, is the most important factor. The lighter the rocket compared to the fuel, the more efficient it will be in term of delta V. As the mass ratio will be fixed by the rocket equation, it also means that the lighter the rocket, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. The engine supports only 3 minutes of operation, will it be enough to reach orbital speed? With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Since a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, an even more powerful engine is required and this is an iterative process. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the wall's thickness and material are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * '''DRY MASS: 12.8552 kg''', WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg '''So here we are. 142kg to carry up to 30km altitude.''' The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten, as well as telemetry equipment. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. Rocket parameters are 15.2 mass ratio (9.0 Delta V, 340 constant I_sp), propellants are LOX and E85 at stoichiometric ratio (no venting, no tank insulation), full aluminium structure. Other parameters are the same than above, and the software used was version 1 (see updates here: [[File:Rocket_mass.c]]). [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> 8caf110db6a383499365730ebd61f428788518c3 6 67 354 2012-11-02T01:41:56Z Vincent 1 Graph representing final acceleration and engine's thrust as function of the initial acceleration (overall thrust-to-weight ratio). wikitext text/x-wiki Graph representing final acceleration and engine's thrust as function of the initial acceleration (overall thrust-to-weight ratio). c3baab58bab6826bf55de9540823e72d2cd4b055 6 68 355 2012-11-02T02:15:49Z Vincent 1 Graph representing rocket dry and wet mass as function of the payload mass (or any other fixed mass subsystem), ranging from 0 to 6 kg payload. wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the payload mass (or any other fixed mass subsystem), ranging from 0 to 6 kg payload. 28ef2f319b623642ed4a320b1b90a914668eade3 6 69 357 2012-11-02T02:47:53Z Vincent 1 Program that iterates to compute an SSTO rocket's dry and wet mass (and other parameters). Compile with gcc -O2 -Wall rocket_mass.c -o rocket_mass -lm for normal use, and add -DMAKE_GRAPH after gcc on the command line if you want to create graph data as wikitext text/x-wiki Program that iterates to compute an SSTO rocket's dry and wet mass (and other parameters). Compile with gcc -O2 -Wall rocket_mass.c -o rocket_mass -lm for normal use, and add -DMAKE_GRAPH after gcc on the command line if you want to create graph data as visible on the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations page]]. 0a2b48fb7fe9475a9377a9fd3f31bac00cb0ae16 372 357 2012-11-07T03:51:53Z Vincent 1 uploaded a new version of "[[File:Rocket mass.c]]":&#32;composite skirts, propellant venting and tank insulation (given data, not computed), mass ratio computed from delta V and Isp. wikitext text/x-wiki Program that iterates to compute an SSTO rocket's dry and wet mass (and other parameters). Compile with gcc -O2 -Wall rocket_mass.c -o rocket_mass -lm for normal use, and add -DMAKE_GRAPH after gcc on the command line if you want to create graph data as visible on the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations page]]. 0a2b48fb7fe9475a9377a9fd3f31bac00cb0ae16 374 372 2012-11-07T18:29:10Z Vincent 1 documentation wikitext text/x-wiki ===Description=== This program is an implementation of the iteration required to compute a rocket's mass. It supports single stage rockets (SSTO) with one engine, takes many input parameters such as thrust-to-weight ratios, propellants data, mass ratio or Isp and Delta V, and computes the dry and wet mass of the rocket. Each subsystem's mass is evaluated and the formulas are simple enough to be changed to suit your needs. ===Features=== See the beginning of the file for all input parameters. Each subsystem has a function computing its mass given the rocket length or other parameters, they can be easily modified. See [[Rocket:First_approximations#Our_first_approximation|first approximation]] for the list of output parameters and an example. The dimensions of the tanks is a major issue. In version 1 and 2, both tanks are computed with the same inner diameter. The goal is to have them roughly the same size so that the rocket can be a simple cylinder. In version 2, tank insulation has been added for the mass concern only. In future version (coming soon) the insulation and wall thickness will be taken into account in the ''outer'' diameter of tanks, so that they can really fit in the same cylinder in any case. Insulation is PU foam, 24kg/m³ density. Tank wall thickness is not properly computed from the pressure and heat, but only estimated as 1mm per 10 atmospheres of pressure, with a minimum value of 1mm for pressurized tanks. The storable fuel tank is currently made of plastic to reduce weight, since pumps will probably be used instead of pressurized tanks. Structural parts are made of aluminium, except the inter-tank skirt (starting from version 2) which is a carbon/epoxy composite. The vehicle's frame is composed of 5 half-square-profiled tubes, 2mm thick, 15mm wide. This ''may'' be enough to support the aircraft staging mechanism and the rocket's weight and flight stress. ===Compilation=== Compile with <code>gcc -O2 -Wall rocket_mass.c -o rocket_mass -lm</code> for normal use, and add <code>-DMAKE_GRAPH</code> after gcc on the command line if you want to create graph data as visible on the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations page]]. da49785d5e672bc4fba6b5e2235c4203b14572ca 376 374 2012-11-07T18:49:48Z Vincent 1 /* Features */ mass ratio stuff wikitext text/x-wiki ===Description=== This program is an implementation of the iteration required to compute a rocket's mass. It supports single stage rockets (SSTO) with one engine, takes many input parameters such as thrust-to-weight ratios, propellants data, mass ratio or Isp and Delta V, and computes the dry and wet mass of the rocket. Each subsystem's mass is evaluated and the formulas are simple enough to be changed to suit your needs. ===Features=== See the beginning of the file for all input parameters. Each subsystem has a function computing its mass given the rocket length or other parameters, they can be easily modified. See [[Rocket:First_approximations#Our_first_approximation|first approximation]] for the list of output parameters and an example. The mass ratio can be either given or computed from I_sp and Delta V. The formula used is very simple but assumes the exhaust velocity (or I_sp) to be constant [https://en.wikipedia.org/wiki/Rocket_equation#Examples (source)]: e^(-DeltaV/V_e), V_e being the gas exhaust velocity, which can be calculated from Isp: V_e = Isp * g₀. The dimensions of the tanks is a major issue. In version 1 and 2, both tanks are computed with the same inner diameter. The goal is to have them roughly the same size so that the rocket can be a simple cylinder. In version 2, tank insulation has been added for the mass concern only. In future version (coming soon) the insulation and wall thickness will be taken into account in the ''outer'' diameter of tanks, so that they can really fit in the same cylinder in any case. Insulation is PU foam, 24kg/m³ density. Tank wall thickness is not properly computed from the pressure and heat, but only estimated as 1mm per 10 atmospheres of pressure, with a minimum value of 1mm for pressurized tanks. The storable fuel tank is currently made of plastic to reduce weight, since pumps will probably be used instead of pressurized tanks. Structural parts are made of aluminium, except the inter-tank skirt (starting from version 2) which is a carbon/epoxy composite. The vehicle's frame is composed of 5 half-square-profiled tubes, 2mm thick, 15mm wide. This ''may'' be enough to support the aircraft staging mechanism and the rocket's weight and flight stress. ===Compilation=== Compile with <code>gcc -O2 -Wall rocket_mass.c -o rocket_mass -lm</code> for normal use, and add <code>-DMAKE_GRAPH</code> after gcc on the command line if you want to create graph data as visible on the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations page]]. 74ffe0ae31a11673467e43ca33e4aa0e9fe16895 388 376 2012-11-09T01:45:54Z Vincent 1 include images for tank model wikitext text/x-wiki ===Description=== This program is an implementation of the iteration required to compute a rocket's mass. It supports single stage rockets (SSTO) with one engine, takes many input parameters such as thrust-to-weight ratios, propellants data, mass ratio or Isp and Delta V, and computes the dry and wet mass of the rocket. Each subsystem's mass is evaluated and the formulas are simple enough to be changed to suit your needs. ===Features=== See the beginning of the file for all input parameters. Each subsystem has a function computing its mass given the rocket length or other parameters, they can be easily modified. See [[Rocket:First_approximations#Our_first_approximation|first approximation]] for the list of output parameters and an example. The mass ratio can be either given or computed from I_sp and Delta V. The formula used is very simple but assumes the exhaust velocity (or I_sp) to be constant [https://en.wikipedia.org/wiki/Rocket_equation#Examples (source)]: e^(-DeltaV/V_e), V_e being the gas exhaust velocity, which can be calculated from Isp: V_e = Isp * g₀. [[File:Rocket_mass_tanks_v1-2.png|right|300px]] The dimensions of the tanks is a major issue. In version 1 and 2, both tanks are computed with the same inner diameter. The goal is to have them roughly the same size so that the rocket can be a simple cylinder. In version 2, tank insulation has been added for the mass concern only. [[File:Rocket_mass_tanks_v3.png|right|300px]] In future version (coming soon) the insulation and wall thickness will be taken into account in the ''outer'' diameter of tanks, so that they can really fit in the same cylinder in any case. Insulation is PU foam, 24kg/m³ density. The insulation thickness will also be calculated from heat and thermal transmission data. Tank wall thickness is not properly computed from the pressure and heat, but only estimated as 1mm per 10 atmospheres of pressure, with a minimum value of 1mm for pressurized tanks. The storable fuel tank is currently made of plastic to reduce weight, since pumps will probably be used instead of pressurized tanks. Structural parts are made of aluminium, except the inter-tank skirt (starting from version 2) which is a carbon/epoxy composite. The vehicle's frame is composed of 5 half-square-profiled tubes, 2mm thick, 15mm wide. This ''may'' be enough to support the aircraft staging mechanism and the rocket's weight and flight stress. ===Compilation=== Compile with <code>gcc -O2 -Wall rocket_mass.c -o rocket_mass -lm</code> for normal use, and add <code>-DMAKE_GRAPH</code> after gcc on the command line if you want to create graph data as visible on the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations page]]. 218bdceee20162a711f30e9553eae1fc2a81d7bd 0 1 359 326 2012-11-04T01:48:43Z Vincent 1 /* News */ nov 2012 news: rocket mass, graphs and twitter wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, you can [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide a low cost orbital launch system for small size satellites, probably with a mass lower than 1kg. Research has led us to consider using an aircraft for rocket launches, the body of the plane being the rocket itself. The rocket, as in any other orbital launch system, would make it to orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''November 2012 update:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 450 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, graphs have been created to represent the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. Look at the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations]] page too. '''''August 2012 update:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may be weakly verified or partly false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design it the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. Can electricity energy be considered for that kind of mission? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, a major issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive and their use is restricted to missiles or UAVs. ===Staging and recovery=== Separation from the rocket is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a complicated autonomous return-to-runway and landing system. ===Guidance=== A satellite navigation system can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. I believe that mere cameras can be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. Anyway, if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: 1) the rocket engine has to be directionally controllable (generally using hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]), or 2) control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) must be used to control the attitude of the rocket, as partially does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but are mandatory in our case. This is one of the big differences between sub-orbital and orbital space flight. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 937f79ac273112a63d8504b0d6154184601955a6 360 359 2012-11-04T02:11:59Z Vincent 1 /* What is the LCAS project? */ project info corrections wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, you can [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''November 2012 update:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 450 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, graphs have been created to represent the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. Look at the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations]] page too. '''''August 2012 update:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may be weakly verified or partly false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with a <i>specific impulse</i> (I_sp), and for rocket engines, it is quite low. However, they are the only engines that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed before falling back on Earth. Rocket trajectories generally tend to form a square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultra-sonic speeds very quickly, the air pressure on their body (mainly the fairing) becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then change trajectory to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around <b>ten times</b> greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some embedded oxidizer. The fact that it uses a turbine design also has a great impact on the improvement of efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since it would be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [http://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, we'll see them below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, around 35km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. This latter point is questionable, since the initial speed of such a plane would still be quite low. Single stage to orbit (SSTO) are also a promising research field for low cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) is the payload. It does not aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design it the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. Can electricity energy be considered for that kind of mission? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, a major issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive and their use is restricted to missiles or UAVs. ===Staging and recovery=== Separation from the rocket is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a complicated autonomous return-to-runway and landing system. ===Guidance=== A satellite navigation system can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. I believe that mere cameras can be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. Anyway, if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: 1) the rocket engine has to be directionally controllable (generally using hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]), or 2) control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) must be used to control the attitude of the rocket, as partially does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but are mandatory in our case. This is one of the big differences between sub-orbital and orbital space flight. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. 8b8c16cf78740a6ba080859a869325494f624c03 361 360 2012-11-04T02:59:40Z Vincent 1 /* How to escape from Earth? */ text fixes wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather [[Founder|my]] researches in the field of astronautics, rocketry and other launch technologies that can be used for the N-Prize competition. It is not an official Web site for the [[N-Prize]]. The official Web site is here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with less than £1000. However, the Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, you can [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''November 2012 update:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 450 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, graphs have been created to represent the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. Look at the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations]] page too. '''''August 2012 update:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may be weakly verified or partly false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a lot of ergols. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircrafts have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, but cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design it the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. Can electricity energy be considered for that kind of mission? If not, what fuel should be used, kerosene, alcohol, E85? Anyway, a major issue with the aircraft is: [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive and their use is restricted to missiles or UAVs. ===Staging and recovery=== Separation from the rocket is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a complicated autonomous return-to-runway and landing system. ===Guidance=== A satellite navigation system can probably be used in the plane for position tracking. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). See the page on the [[EmbeddedRocketComputer|embedded computer]]. ==The rocket== Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for examples. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. The loss of weight due to that water is often a good thing because it burns producing so much heat that the water can keep the engine cool enough to survive. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. To my eyes, alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is only 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? I believe that '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good (regenerative) cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== Liquid Oxygen (LOX) is the obvious/best choice for high Isp. However, it has lots of drawbacks because of the need for cryogenics storage, cautious manipulation, and engine design, that make it quite expensive and much complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Other leads should then be explored, like [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide]. [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide] would even be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. More details on the [[RocketEngines|rocket engines page]]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. I believe that mere cameras can be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. Anyway, if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: 1) the rocket engine has to be directionally controllable (generally using hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]), or 2) control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) must be used to control the attitude of the rocket, as partially does SpaceX with the Merlin engine. Both cases imply complications on the rocket's and engine hardware, but are mandatory in our case. This is one of the big differences between sub-orbital and orbital space flight. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing device, radioactive, EM emitting, mirrors... The ground segment will have to be developed from scratch since I don't think anybody would mind tracking 20g 100miles away. d7f485e6a586eb950324cd3443cd7cf35efaa407 362 361 2012-11-05T16:33:09Z Vincent 1 /* N-Prize and reflections on low-cost access to space */ text fixes and adding main pages links wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''November 2012 update:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 450 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, graphs have been created to represent the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. Look at the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations]] page too. '''''August 2012 update:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may be weakly verified or partly false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen (LOX) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics storage, cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. baad71c17ef340a29bf27cd225247d58194887be 377 362 2012-11-08T02:08:20Z Vincent 1 /* Oxidizer */ links wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''November 2012 update:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 450 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, graphs have been created to represent the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. Look at the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations]] page too. '''''August 2012 update:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may be weakly verified or partly false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. f562fb877bb0a6cf0ccf53b7a489a50131cba0e5 398 377 2012-11-11T04:29:45Z Vincent 1 /* The aircraft */ links wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''November 2012 update:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 450 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, graphs have been created to represent the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. Look at the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations]] page too. '''''August 2012 update:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may be weakly verified or partly false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. 357b30ae79ff9959725b4346a58953a580287732 0 70 363 2012-11-05T19:23:19Z Vincent 1 page creation wikitext text/x-wiki =Aircraft operation= This page is related to aircraft lift-off, staging and landing operations. The return capabilities are discussed too. ==Lift-off== Lift-off will not be simple. The engines we are designing are optimized for high-altitude flight, and may not be able to provide the required thrust at ground level. Some additional acceleration mechanisms would then be required: * using the rocket's engine for a short time. The problem is then how to carry the additional propellant required for this lift-off operation. An additional disposable tank could be mounted on the rocket or on the aircraft, jettisoned after this burn. The engine may allow throttling, reducing the thrust and propellant consumption, while still providing the required delta V for lift-off. * push-on boosters, either with propellers or miniature solid rocket engines, that would stay on the ground when the plane lifts-off. * some kind of pulling cable on the runway ==Staging and risk evaluation== In the subsonic flight variant, staging will occur when the highest service altitude has been reached, which will probably happen when the fuel has all been used. Possible staging mechanism: * aircraft is attached to the rocket by some pins, the rocket is released then fired with no possible abort. A self-destruct mechanism must be added to the rocket, because if the rocket engine fails to ignite, the rocket would turn into a ballistic air-to-surface missile. * aircraft is attached to the rocket by some pins, the rocket is fired before being released. Risks are lower, but the attachment may be complicated to design, since the thrust of the rocket would be acting on it. * rocket is attached to the aircraft with a rail on which it can roll. The rocket engine would ignite before separating from the aircraft, providing a way to return to the ground without exploding if it doesn't ignite, by gliding. The rocket engine's thrust is what would make the rocket separate, which is probably the safest and simplest way to make it. The self-destruct could be required in all cases, because there are lots of other issues than separation. ==Return to earth mission== What should be done with the "aircraft" when it has separated with the rocket? What remains is mainly the wings and the tail, probably linked together by spars. For the N-prize, the aircraft should be reusable to deduce its cost from the project's cost. For other applications, this is not mandatory. The engines could be complicated and expensive to build, to it may be important to return them safely to earth, if they have a lifespan of at least two flight durations. With two engines mounted under each wing, the returning aircraft would be hard to land properly. There is '''no landing gear''', and the engines are the lowest part of the aircraft. If it is placed on the floor, the engines and the tail support the wings. A landing gear could consist of wheels deeply embedded in the engines and in the tail for its simplest form, or of a deployable gear stored in the wings or in the attachment between the wings and the tail. List of strategies for aircraft return: * free fall: ** crash on ground (no safe return) ** parachute deployment triggered by the altitude or by timer after separation (safe return is not guaranteed for aircraft, and unlikely if the rocket has not separated) ** airbag, [http://marsrovers.jpl.nasa.gov MER] style. If used without parachute, the wings at least will be damaged (no safe return in all case). If used in combination with the parachute system, airbags could prevent the engines from being destroyed at touchdown (safe return for aicraft, no safe return for non-separated rocket). * controlled gliding, '''requires''' the centre of mass to allow it and probably some balancing weights to have it at the right position; '''requires''' and independent battery to operate the plane actuators for the return trip. The landing is then the issue: ** controlled crash at zero vertical speed, like a miniature remote-controlled aircraft, without landing gear. If autonomous, requires to be lucky with the return terrain, or to be able to return to the runway with unpowered flight, possibly by recording wing directions on the way up. With no landing gear, landing would be on the engines (no safe return of the vehicle). If remote-controlled, the terrain may be better. ** landing at zero vertical speed, with a landing gear. It still depends on the landing terrain and the ability to find and go to a runway, but engine could be spared (safe return possible). If remote controlled, it depends on the ability to fly to glide to something close enough to a runway. ** controlled crash, upside down. Having the aircraft upside down may not damage the engines, but damage the tail instead. This could be complicated to do autonomously, and would still depend on the landing terrain (safe engine return possible). * powered flight could be used to return the complete vehicle to a runway and allow for a safe return on non catastrophic rocket failure and no aircraft failure. If autonomous, that '''requires''' a complicated control system to return to runway. If remote-controlled, that looks more easy and could be the safest solution for abort or safe return. Requires one of the following propulsion systems: ** continuing turbofan operation, '''requires''' more fuel in the aircraft ** electric motor with a propeller, '''requires''' additional battery If the aircraft has to return to ground autonomously, a control card has to be put in it, managing several sensors and actuators. Some of the sensors would also have to be in the rocket. This redundancy could allow the vehicle to return to the ground safely if something wrong happens with the rocket's control system. The remote-controlled operation will probably be developed for early tests of the aircraft prototype, and can thus be left in the final vehicle for return and especially landing. If recovery is required, a beacon should be put aboard the aircraft. 49e58f4a90f13f67e879ebf38f63fb99a32b4274 0 71 364 2012-11-05T21:51:52Z Vincent 1 starting propellants wikitext text/x-wiki =Propellants= ==Fuels== * RP-1, storable, not dangerous, available?, moderately cheap, density: 806 kg/m3 * E85, storable, not dangerous, readily available, cheap, density: 780 kg/m3 * H2, cryogenic, not dangerous, requires specific storage and permit?, not cheap?, density: 71 kg/m3 * ETHANOL, storable, not dangerous, readily available, cheap, 92.5% density: 800 kg/m3 ==Oxidizers== * [https://en.wikipedia.org/wiki/Liquid_oxygen LOX], cryogenic, explosive, requires cryogenic storage, cheap, density: 1141 kg/m3 at 92.2K and 1 atm, 974.42 kg/m3 at 120K and 10bar * [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] (N₂O), refrigerated liquid (boiling at -88.5°C), non-toxic, density: 1223 kg/m3 * [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide], pressurized, dangerous, expensive to have it manufactured at a high concentration, density: 1450 kg/m3 (pure) ==Mixes== {| border="1" class="wikitable" !Oxidizer !Fuel !I_sp (ground) !I_sp (vacuum) !Stoichiometric !Combustion temp. !Average density |- |rowspan="4"|LOX |LH₂ |? |455 |8 |? |? |- |RP-1 |? |358 |2.56 |? |? |- |E85 |? |? |2.26 |? |? |- |Ethanol 92.5% |? |? |2.19 |? |? |} 6147e9fe1b9439ced7434298552c155c1349f876 365 364 2012-11-06T00:17:38Z Vincent 1 propellants basic data wikitext text/x-wiki =Propellants= Here are some links: [http://www.thespacerace.com/forum/index.php?topic=2583.0 tables and info]. Wikipedia's [https://en.wikipedia.org/wiki/Liquid_rocket_propellants#Bipropellants table]. Another [http://www.braeunig.us/space/propel.htm table]. ==Fuels== * RP-1, storable, not dangerous, available?, moderately cheap, density: 806 kg/m³ * E85, storable, not dangerous, readily available, cheap, density: 780 kg/m³ * LH₂, cryogenic, not dangerous, requires specific storage and permit?, quite expensive, density: 71 kg/m³ * Ethanol, storable, not dangerous, readily available, cheap, density at 92.5%: 800 kg/m³ ==Oxidizers== Besides the cryogenic issue, LOX is probably the safest oxidizer. Others may be storable at nearly ambient temperature, or under pressure, but they are less stable, subject to explosion or toxic, and more expensive. In a cold country, Nitrous oxide can be a good alternative, but if temperature is around 20°C its density is too low and tanks require a large spare volume. Nitrous oxide is at least 20 times more expensive than LOX too. * [https://en.wikipedia.org/wiki/Liquid_oxygen LOX], cryogenic, explosive, requires cryogenic storage, cheap, density: 1141 kg/m³ at 92.2K and 1 atm, 974.42 kg/m³ at 120K and 10bar * [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] (N₂O), refrigerated liquid (boiling at -88.5°C) or self pressurizing (vapour pressure at 20°C is ~50.1 bar), but critical temperature is at 36.4°C, non-toxic, quite expensive, density: 1223 kg/m³ at -88.5°C, 750 kg/m³ at 20°C, changes dramatically with temperature * [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide], pressurized, self-decomposes explosively, expensive to have it manufactured at a high concentration, density: 1450 kg/m³ (pure) * [https://en.wikipedia.org/wiki/Nitrogen_tetroxide Nitrogen tetroxide] (N₂O₄), storable, highly toxic, density: 1443 kg/m³ at 21°C. ==Mixes== The stoichiometric ratio is not the optimal ratio for rocket applications. The density and exhaust speeds are more important than maximum temperature. In particular, the ratio T_c/M, combustion temperature / molecular mass, is a good indicator of the exhaust speed, as explained [http://www.thespacerace.com/forum/index.php?topic=2583.msg17485#msg17485 here]. {| border="1" class="wikitable" !Oxidizer !Fuel !I_sp (sea lvl) !max I_sp (vacuum) !Stoichiometric !T_c Combustion temp. (K) !Average density (kg/l) |- |rowspan="4"|LOX |LH₂ |381 (r=5.0) |455 |8 |3304 |0.32 |- |RP-1 |289 (r=2.29) |353 |2.56 |3526 |1.02 |- |E85 |? |? |2.26 |around 3360 |around 1 |- |Ethanol 95% |277 |? |2.19 |3314 |0.97 |} 8fec021d0b7fdcb9973f14a6e98d78614eafe8b8 368 365 2012-11-06T03:54:56Z Vincent 1 /* Oxidizers */ LOX page link wikitext text/x-wiki =Propellants= Here are some links: [http://www.thespacerace.com/forum/index.php?topic=2583.0 tables and info]. Wikipedia's [https://en.wikipedia.org/wiki/Liquid_rocket_propellants#Bipropellants table]. Another [http://www.braeunig.us/space/propel.htm table]. ==Fuels== * RP-1, storable, not dangerous, available?, moderately cheap, density: 806 kg/m³ * E85, storable, not dangerous, readily available, cheap, density: 780 kg/m³ * LH₂, cryogenic, not dangerous, requires specific storage and permit?, quite expensive, density: 71 kg/m³ * Ethanol, storable, not dangerous, readily available, cheap, density at 92.5%: 800 kg/m³ ==Oxidizers== Besides the cryogenic issue, LOX is probably the safest oxidizer. Others may be storable at nearly ambient temperature, or under pressure, but they are less stable, subject to explosion or toxic, and more expensive. In a cold country, Nitrous oxide can be a good alternative, but if temperature is around 20°C its density is too low and tanks require a large spare volume. Nitrous oxide is at least 20 times more expensive than LOX too. * [https://en.wikipedia.org/wiki/Liquid_oxygen LOX] ([[LOX|internal link]]), cryogenic, explosive, requires cryogenic storage, cheap, density: 1141 kg/m³ at 92.2K and 1 atm, 974.42 kg/m³ at 120K and 10bar * [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] (N₂O), refrigerated liquid (boiling at -88.5°C) or self pressurizing (vapour pressure at 20°C is ~50.1 bar), but critical temperature is at 36.4°C, non-toxic, quite expensive, density: 1223 kg/m³ at -88.5°C, 750 kg/m³ at 20°C, changes dramatically with temperature * [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide], pressurized, self-decomposes explosively, expensive to have it manufactured at a high concentration, density: 1450 kg/m³ (pure) * [https://en.wikipedia.org/wiki/Nitrogen_tetroxide Nitrogen tetroxide] (N₂O₄), storable, highly toxic, density: 1443 kg/m³ at 21°C. ==Mixes== The stoichiometric ratio is not the optimal ratio for rocket applications. The density and exhaust speeds are more important than maximum temperature. In particular, the ratio T_c/M, combustion temperature / molecular mass, is a good indicator of the exhaust speed, as explained [http://www.thespacerace.com/forum/index.php?topic=2583.msg17485#msg17485 here]. {| border="1" class="wikitable" !Oxidizer !Fuel !I_sp (sea lvl) !max I_sp (vacuum) !Stoichiometric !T_c Combustion temp. (K) !Average density (kg/l) |- |rowspan="4"|LOX |LH₂ |381 (r=5.0) |455 |8 |3304 |0.32 |- |RP-1 |289 (r=2.29) |353 |2.56 |3526 |1.02 |- |E85 |? |? |2.26 |around 3360 |around 1 |- |Ethanol 95% |277 |? |2.19 |3314 |0.97 |} 8cacbe15bf527b4e7d2f9a098e35dd6fc9f5902a 379 368 2012-11-08T02:47:34Z Vincent 1 /* Oxidizers */ nox critical point wikitext text/x-wiki =Propellants= Here are some links: [http://www.thespacerace.com/forum/index.php?topic=2583.0 tables and info]. Wikipedia's [https://en.wikipedia.org/wiki/Liquid_rocket_propellants#Bipropellants table]. Another [http://www.braeunig.us/space/propel.htm table]. ==Fuels== * RP-1, storable, not dangerous, available?, moderately cheap, density: 806 kg/m³ * E85, storable, not dangerous, readily available, cheap, density: 780 kg/m³ * LH₂, cryogenic, not dangerous, requires specific storage and permit?, quite expensive, density: 71 kg/m³ * Ethanol, storable, not dangerous, readily available, cheap, density at 92.5%: 800 kg/m³ ==Oxidizers== Besides the cryogenic issue, LOX is probably the safest oxidizer. Others may be storable at nearly ambient temperature, or under pressure, but they are less stable, subject to explosion or toxic, and more expensive. In a cold country, Nitrous oxide can be a good alternative, but if temperature is around 20°C its density is too low and tanks require a large spare volume. Nitrous oxide is at least 20 times more expensive than LOX too. * [https://en.wikipedia.org/wiki/Liquid_oxygen LOX] ([[LOX|internal link]]), cryogenic, explosive, requires cryogenic storage, cheap, density: 1141 kg/m³ at 92.2K and 1 atm, 974.42 kg/m³ at 120K and 10bar * [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] (N₂O), refrigerated liquid (boiling at -88.5°C) or self pressurizing (vapour pressure at 20°C is ~50.1 bar), but critical point is 36.4°C and 72.45 bar, non-toxic, quite expensive, density: 1223 kg/m³ at -88.5°C, 750 kg/m³ at 20°C, changes dramatically with temperature * [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide], pressurized, self-decomposes explosively, expensive to have it manufactured at a high concentration, density: 1450 kg/m³ (pure) * [https://en.wikipedia.org/wiki/Nitrogen_tetroxide Nitrogen tetroxide] (N₂O₄), storable, highly toxic, density: 1443 kg/m³ at 21°C. ==Mixes== The stoichiometric ratio is not the optimal ratio for rocket applications. The density and exhaust speeds are more important than maximum temperature. In particular, the ratio T_c/M, combustion temperature / molecular mass, is a good indicator of the exhaust speed, as explained [http://www.thespacerace.com/forum/index.php?topic=2583.msg17485#msg17485 here]. {| border="1" class="wikitable" !Oxidizer !Fuel !I_sp (sea lvl) !max I_sp (vacuum) !Stoichiometric !T_c Combustion temp. (K) !Average density (kg/l) |- |rowspan="4"|LOX |LH₂ |381 (r=5.0) |455 |8 |3304 |0.32 |- |RP-1 |289 (r=2.29) |353 |2.56 |3526 |1.02 |- |E85 |? |? |2.26 |around 3360 |around 1 |- |Ethanol 95% |277 |? |2.19 |3314 |0.97 |} fb708cc77ec4db686fbf43bbe14d07710dc58ef1 0 62 367 344 2012-11-06T03:54:26Z Vincent 1 /* Liquid Oxygen */ propellants link wikitext text/x-wiki =Liquid Oxygen= See [[Propellants]] page and [https://en.wikipedia.org/wiki/Oxygen Wikipedia's page]. Boiling point of -182.95°C (90.20 K). Critical temperature of -118.59°C (154.59 K) meaning that it cannot stay liquid above this temperature. 1 Litre of LOX provides 840 Litre of GOX (gaseous oxygen) at 1atm/20°C. Like any liquid, LOX vaporizes when stored above its boiling point. The vapour builds up to pressure that the tank will no be able to contain, so a relief valve must be put in place to compensate this evaporation. Insulation limits the rate of vaporization, the best being the vacuum space between a tank inner and outer walls, like Dewar flasks. 758223b24f176dec44ece4baa443f7119c34abb0 373 367 2012-11-07T03:57:13Z Vincent 1 /* Liquid Oxygen */ latent heat of vaporization wikitext text/x-wiki =Liquid Oxygen= See [[Propellants]] page and [https://en.wikipedia.org/wiki/Oxygen Wikipedia's page]. Boiling point of -182.95°C (90.20 K). Critical temperature of -118.59°C (154.59 K) meaning that it cannot stay liquid above this temperature. Density: 141 kg/m3 at 92.2K and 1 atm, 974.42 kg/m3 at 120K and 10bar 1 Litre of LOX provides 840 Litre of GOX (gaseous oxygen) at 1atm/20°C. Like any liquid, LOX vaporizes when stored above its boiling point. The vapour builds up to pressure that the tank will no be able to contain, so a relief valve must be put in place to compensate this evaporation. Insulation limits the rate of vaporization, the best being the vacuum space between a tank inner and outer walls, like Dewar flasks. Latent heat of vaporization: 213 kJ/kg. 6d17f7bd47ddb117de609d17573fc1d9afd75d8d 378 373 2012-11-08T02:44:20Z Vincent 1 /* Liquid Oxygen */ fixes wikitext text/x-wiki =Liquid Oxygen= See [[Propellants]] page and [https://en.wikipedia.org/wiki/Oxygen Wikipedia's page]. Boiling point of -182.95°C (90.20 K). Critical temperature of -118.59°C (154.59 K) meaning that it cannot stay liquid above this temperature. Critical pressure is 5.043 MPa (49.77 atm). Density: 1141 kg/m3 at 92.2K and 1 atm, 974.42 kg/m3 at 120K and 10bar 1 Litre of LOX provides 840 Litre of GOX (gaseous oxygen) at 1atm/20°C. Like any liquid, LOX vaporizes when stored above its boiling point. The vapour builds up to pressure that the tank will no be able to contain, so a relief valve must be put in place to compensate this evaporation. Insulation limits the rate of vaporization, the best being the vacuum space between a tank inner and outer walls, like Dewar flasks. Latent heat of vaporization: 213 kJ/kg (6.82 kJ/mol). c1390a40c20731eaa8e3fc49e0a47323effb2d2a 383 378 2012-11-08T18:00:54Z Vincent 1 /* Liquid Oxygen */ tank link wikitext text/x-wiki =Liquid Oxygen= ''See [[Propellants]] page and [https://en.wikipedia.org/wiki/Oxygen Wikipedia's page]. See also [[Rocket_Main_Tank#Cryogenic_fuel_tanks|tank page]] for insulation.'' Boiling point of -182.95°C (90.20 K). Critical temperature of -118.59°C (154.59 K) meaning that it cannot stay liquid above this temperature. Critical pressure is 5.043 MPa (49.77 atm). Density: 1141 kg/m3 at 92.2K and 1 atm, 974.42 kg/m3 at 120K and 10bar 1 Litre of LOX provides 840 Litre of GOX (gaseous oxygen) at 1atm/20°C. Like any liquid, LOX vaporizes when stored above its boiling point. The vapour builds up to pressure that the tank will no be able to contain, so a relief valve must be put in place to compensate this evaporation. Insulation limits the rate of vaporization, the best being the vacuum space between a tank inner and outer walls, like Dewar flasks. Latent heat of vaporization: 213 kJ/kg (6.82 kJ/mol). e512c57af4e980d793d3ecd175d382892ea1db47 0 24 369 115 2012-11-06T04:37:30Z Vincent 1 more things to say about tanks wikitext text/x-wiki =Rocket Fuel tanks= In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. [[Image:Restone_tank.jpg|center|250px]] ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Cryogenic fuel tanks== ==Wall thickness and material== Tank material mostly depend on money and on what's available on the market. For robust tanks, we will use 6061 Aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. Aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. ==Accessories== Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. ==Thermal insulation== Cryogenic fuel tanks may require to be insulated to limit vaporization or even boiling. A list of thermal conductivities is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. Expanded polystyrene has a thermal conductivity of 0.03 W/m.K, PU foam has a 0.02, mineral insulation 0.04, neoprene 0.054, cotton 0.03. A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. b6ca91b37656c12894cc97da970190d417418459 370 369 2012-11-07T03:04:04Z Vincent 1 /* Thermal insulation */ table and densities wikitext text/x-wiki =Rocket Fuel tanks= In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. [[Image:Restone_tank.jpg|center|250px]] ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Cryogenic fuel tanks== ==Wall thickness and material== Tank material mostly depend on money and on what's available on the market. For robust tanks, we will use 6061 Aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. Aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. ==Accessories== Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. ==Thermal insulation== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivities] (''k'') is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayedq |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | | |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. d6dba98e1e8309b549a5c7bcd68f288e968f876d 380 370 2012-11-08T03:12:16Z Vincent 1 /* Cryogenic fuel tanks */ about evaporation wikitext text/x-wiki =Rocket Fuel tanks= In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. [[Image:Restone_tank.jpg|center|250px]] ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Cryogenic fuel tanks== ''See [[LOX]] page too.'' Cryogenic and also low boiling temperature liquids like nitrous oxide are persistently evaporating at ambient temperature. It's like having water at 100°C and providing always more heat to it. When the vapour pressure is high enough, and when tanks are solid enough too, the evaporation can reach equilibrium and the tank can contain a stable mix of liquid and gas at high pressure. This is the case for nitrous oxide at temperatures below 36.4°C, its critical temperature above which it turns all into gas, no matter what pressure is used. The issue then becomes the density of the mixture, which drops greatly. For LOX, the critical temperature is -118.59°C, and the critical pressure is 50.43 bar. There's no point in keeping it so much pressurized because it could just boil off at this temperature. Since the phase change occurs at a constant temperature, we can as well choose a temperature and a pressure at which the LOX density is high enough, but that's a trade-off with the evaporation rate. Since the temperature difference between inside and outside the tank is greater, even more heat is transferred to the LOX, and more evaporation is created. Tank insulation is then required to avoid venting all the propellant before actually using it (balloon or aircraft launch can take some time to get to the launch altitude). ==Wall thickness and material== Tank material mostly depend on money and on what's available on the market. For robust tanks, we will use 6061 Aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. Aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. ==Accessories== Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. ==Thermal insulation== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivities] (''k'') is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayedq |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | | |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. fed40635c47741cc02c062b2d5390812abed1b2e 381 380 2012-11-08T04:13:20Z Vincent 1 /* Wall thickness and material */ material compat. wikitext text/x-wiki =Rocket Fuel tanks= In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. [[Image:Restone_tank.jpg|center|250px]] ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Cryogenic fuel tanks== ''See [[LOX]] page too.'' Cryogenic and also low boiling temperature liquids like nitrous oxide are persistently evaporating at ambient temperature. It's like having water at 100°C and providing always more heat to it. When the vapour pressure is high enough, and when tanks are solid enough too, the evaporation can reach equilibrium and the tank can contain a stable mix of liquid and gas at high pressure. This is the case for nitrous oxide at temperatures below 36.4°C, its critical temperature above which it turns all into gas, no matter what pressure is used. The issue then becomes the density of the mixture, which drops greatly. For LOX, the critical temperature is -118.59°C, and the critical pressure is 50.43 bar. There's no point in keeping it so much pressurized because it could just boil off at this temperature. Since the phase change occurs at a constant temperature, we can as well choose a temperature and a pressure at which the LOX density is high enough, but that's a trade-off with the evaporation rate. Since the temperature difference between inside and outside the tank is greater, even more heat is transferred to the LOX, and more evaporation is created. Tank insulation is then required to avoid venting all the propellant before actually using it (balloon or aircraft launch can take some time to get to the launch altitude). ==Wall thickness and material== Tank material first has to be stable with what's inside. Lists are available for cryogenic liquids at least. Besides this basic filter, the material choice mostly depends on money and on what's available on the market. For pressurized tanks, we will use 6061 aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. Aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. ==Accessories== Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. ==Thermal insulation== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivities] (''k'') is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayedq |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | | |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. 7d909340591f194aa15775a453b248af20e6c0bb 382 381 2012-11-08T04:13:54Z Vincent 1 /* Thermal insulation */ typo wikitext text/x-wiki =Rocket Fuel tanks= In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. [[Image:Restone_tank.jpg|center|250px]] ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Cryogenic fuel tanks== ''See [[LOX]] page too.'' Cryogenic and also low boiling temperature liquids like nitrous oxide are persistently evaporating at ambient temperature. It's like having water at 100°C and providing always more heat to it. When the vapour pressure is high enough, and when tanks are solid enough too, the evaporation can reach equilibrium and the tank can contain a stable mix of liquid and gas at high pressure. This is the case for nitrous oxide at temperatures below 36.4°C, its critical temperature above which it turns all into gas, no matter what pressure is used. The issue then becomes the density of the mixture, which drops greatly. For LOX, the critical temperature is -118.59°C, and the critical pressure is 50.43 bar. There's no point in keeping it so much pressurized because it could just boil off at this temperature. Since the phase change occurs at a constant temperature, we can as well choose a temperature and a pressure at which the LOX density is high enough, but that's a trade-off with the evaporation rate. Since the temperature difference between inside and outside the tank is greater, even more heat is transferred to the LOX, and more evaporation is created. Tank insulation is then required to avoid venting all the propellant before actually using it (balloon or aircraft launch can take some time to get to the launch altitude). ==Wall thickness and material== Tank material first has to be stable with what's inside. Lists are available for cryogenic liquids at least. Besides this basic filter, the material choice mostly depends on money and on what's available on the market. For pressurized tanks, we will use 6061 aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. Aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. ==Accessories== Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. ==Thermal insulation== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivities] (''k'') is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayed |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | | |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. a89eda9132c41e4fa9098a557b11e53fadfb41a1 384 382 2012-11-08T23:46:40Z Vincent 1 page refactoring, more about insulation heat transmission and single divided tank wikitext text/x-wiki =Rocket Fuel tanks= [[Image:Restone_tank.jpg|right|320px]] In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. The sphere is the most lightweight volume (volume / area is minimized), but we can't have rockets as large as they are long, so cylinders with hemispheric caps are used. Having a single tank cut in two like for the Restone is efficient mass-wise and volume-wise but can bring new issues for insulation in case of a single cryogenic fluid (funny enough, that was the case for the Restone which used LOX and ethanol). The thicker insulation may overtake the mass benefits of a single tank. ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Wall thickness and material== Tank material first has to be stable with what's inside. Lists are available for cryogenic liquids at least. Besides this basic filter, the material choice mostly depends on money and on what's available on the market. For pressurized tanks, we will use 6061 aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. For example, aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. For a 0.2m wide tank, the thickness can be 1mm for the same pressure, or twice the pressure for the same thickness. ==Cryogenic fuel tanks== Cryogenic and also low boiling temperature liquids like nitrous oxide are persistently evaporating at ambient temperature. It's like having water at 100°C and providing always more heat to it. When the vapour pressure is high enough, and when tanks are solid enough too, the evaporation can reach equilibrium and the tank can contain a stable mix of liquid and gas at high pressure. This is the case for nitrous oxide at temperatures below 36.4°C, its critical temperature above which it turns all into gas, no matter what pressure is used. The issue then becomes the density of the mixture, which drops greatly. For [[LOX]], the critical temperature is -118.59°C, and the critical pressure is 50.43 bar. There's no point in keeping it so much pressurized because it could just boil off at this temperature. Since the phase change occurs at a constant temperature, we can as well choose a temperature and a pressure at which the LOX density is high enough, but that's a trade-off with the evaporation rate. Since the temperature difference between inside and outside the tank is greater, even more heat is transferred to the LOX, and more evaporation is created. Tank insulation is then required to avoid venting all the propellant before actually using it (balloon or aircraft launch can take some time to get to the launch altitude). A material has a [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K), representing its ability to conduct heat. An insulation layer has a [https://en.wikipedia.org/wiki/R-value_(insulation) thermal resistance] (R-value) and its opposite, the [https://en.wikipedia.org/wiki/U-value#U-value thermal transmittance] (U-value), indicating how much resistance to heat the material provides. For an insulation layer of thickness ''L'', ''R = L/k'' and ''U = k/L''. ===Thermal insulation materials=== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of thermal conductivities is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayed |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | |readily available, cheap |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. ==Propellant lines== Pumps and Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. 761f458642cba6e2d56a90a68abcf359fc5c2690 389 384 2012-11-09T04:51:41Z Vincent 1 /* Cryogenic fuel tanks */ evaporation rate wikitext text/x-wiki =Rocket Fuel tanks= [[Image:Restone_tank.jpg|right|320px]] In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. The sphere is the most lightweight volume (volume / area is minimized), but we can't have rockets as large as they are long, so cylinders with hemispheric caps are used. Having a single tank cut in two like for the Restone is efficient mass-wise and volume-wise but can bring new issues for insulation in case of a single cryogenic fluid (funny enough, that was the case for the Restone which used LOX and ethanol). The thicker insulation may overtake the mass benefits of a single tank. ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Wall thickness and material== Tank material first has to be stable with what's inside. Lists are available for cryogenic liquids at least. Besides this basic filter, the material choice mostly depends on money and on what's available on the market. For pressurized tanks, we will use 6061 aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. For example, aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. For a 0.2m wide tank, the thickness can be 1mm for the same pressure, or twice the pressure for the same thickness. ==Cryogenic fuel tanks== Cryogenic and also low boiling temperature liquids like nitrous oxide are persistently evaporating at ambient temperature. It's like having water at 100°C and providing always more heat to it. When the vapour pressure is high enough, and when tanks are solid enough too, the evaporation can reach [https://en.wikipedia.org/wiki/Evaporation#Evaporative_equilibrium equilibrium] and the tank can contain a stable mix of liquid and gas at high pressure. This is the case for nitrous oxide at temperatures below 36.4°C, its critical temperature above which it turns all into gas, no matter what pressure is used. The issue then becomes the density of the mixture, which drops greatly. For [[LOX]], the critical temperature is -118.59°C, and the critical pressure is 50.43 bar. There's no point in keeping it so much pressurized because it could just boil off at this temperature. Since the phase change occurs at a constant temperature, we can as well choose a temperature and a pressure at which the LOX density is high enough, but that's a trade-off with the evaporation rate. Since the temperature difference between inside and outside the tank is greater, even more heat is transferred to the LOX, and more evaporation is created. Tank insulation is then required to avoid venting all the propellant before actually using it (balloon or aircraft launch can take some time to get to the launch altitude). ===Calculating evaporation rate=== A material has a [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K), representing its ability to conduct heat. An insulation layer has a [https://en.wikipedia.org/wiki/R-value_(insulation) thermal resistance] (R-value) and its opposite, the [https://en.wikipedia.org/wiki/U-value#U-value thermal transmittance] (U-value), indicating how much resistance to heat the material provides. For an insulation layer of thickness ''L'', ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The heat transfer by convection is then given by the formula: Φ = A × U × (T1 - T2), in Watt. The thermal transmittance of the tank material can be ignored for first approximations. A is the area where the heat exchange is made, the outer surface of the insulation layer. T1 is the convecting air temperature, T2 is external tank temperature, which is close enough from the internal temperature (constant, boiling point given by pressure) to be taken for it. Examples can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc]. Finally, the evaporation rate is the heat of vaporization ΔH_vap/Φ in kg/s. That requires to know ΔH_vap for the chosen storage temperature, but graphs are available for common molecules like O2. ===Thermal insulation materials=== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of thermal conductivities is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayed |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | |readily available, cheap |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. ==Propellant lines== Pumps and Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. b8906c401bc518b872d9ec682e0a27aa8a6fec8c 390 389 2012-11-09T14:47:51Z Vincent 1 aerogel wikitext text/x-wiki =Rocket Fuel tanks= [[Image:Restone_tank.jpg|right|320px]] In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. The sphere is the most lightweight volume (volume / area is minimized), but we can't have rockets as large as they are long, so cylinders with hemispheric caps are used. Having a single tank cut in two like for the Restone is efficient mass-wise and volume-wise but can bring new issues for insulation in case of a single cryogenic fluid (funny enough, that was the case for the Restone which used LOX and ethanol). The thicker insulation may overtake the mass benefits of a single tank. ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Wall thickness and material== Tank material first has to be stable with what's inside. Lists are available for cryogenic liquids at least. Besides this basic filter, the material choice mostly depends on money and on what's available on the market. For pressurized tanks, we will use 6061 aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. For example, aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. For a 0.2m wide tank, the thickness can be 1mm for the same pressure, or twice the pressure for the same thickness. ==Cryogenic fuel tanks== Cryogenic and also low boiling temperature liquids like nitrous oxide are persistently evaporating at ambient temperature. It's like having water at 100°C and providing always more heat to it. When the vapour pressure is high enough, and when tanks are solid enough too, the evaporation can reach [https://en.wikipedia.org/wiki/Evaporation#Evaporative_equilibrium equilibrium] and the tank can contain a stable mix of liquid and gas at high pressure. This is the case for nitrous oxide at temperatures below 36.4°C, its critical temperature above which it turns all into gas, no matter what pressure is used. The issue then becomes the density of the mixture, which drops greatly. For [[LOX]], the critical temperature is -118.59°C, and the critical pressure is 50.43 bar. There's no point in keeping it so much pressurized because it could just boil off at this temperature. Since the phase change occurs at a constant temperature, we can as well choose a temperature and a pressure at which the LOX density is high enough, but that's a trade-off with the evaporation rate. Since the temperature difference between inside and outside the tank is greater, even more heat is transferred to the LOX, and more evaporation is created. Tank insulation is then required to avoid venting all the propellant before actually using it (balloon or aircraft launch can take some time to get to the launch altitude). ===Calculating evaporation rate=== A material has a [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K), representing its ability to conduct heat. An insulation layer has a [https://en.wikipedia.org/wiki/R-value_(insulation) thermal resistance] (R-value) and its opposite, the [https://en.wikipedia.org/wiki/U-value#U-value thermal transmittance] (U-value), indicating how much resistance to heat the material provides. For an insulation layer of thickness ''L'', ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The heat transfer by convection is then given by the formula: Φ = A × U × (T1 - T2), in Watt. The thermal transmittance of the tank material can be ignored for first approximations. A is the area where the heat exchange is made, the outer surface of the insulation layer. T1 is the convecting air temperature, T2 is external tank temperature, which is close enough from the internal temperature (constant, boiling point given by pressure) to be taken for it. Examples can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc]. Finally, the evaporation rate is the heat of vaporization ΔH_vap/Φ in kg/s. That requires to know ΔH_vap for the chosen storage temperature, but graphs are available for common molecules like O2. ===Thermal insulation materials=== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of thermal conductivities is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. PU foam is a simple solution, it can be sprayed at the desired thickness but may be hard to spray in very thin layers. Aerogel is the best existing insulation material and it has been used by Armadillo Aerospace (see [[Rocket_Main_Tank#Resources|below]]). Aerogel for cryogenic applications is [http://www.aerogel.com/products/overview-product.html currently available] in 5mm or 10mm thick sheets. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |Aerogel |15 |13 |readily available in sheets, cheap |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayed |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | |readily available, cheap |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. ==Propellant lines== Pumps and Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. ==Resources== Armadillo Aerospace has [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=333 documented] their LOX tank insulation: We settled on using Pyrogel insulating blankets from Aspen Aerogels to insulate our tanks: [http://www.aerogel.com/products/pdf/Pyrogel_6250_DS.pdf#search=%22pyrogel%20insulation%22] After Phil figured out the right pattern to make the gores for the spherical tanks, it went very well. We used a spray adhesive to attach it, and we tested all the combinations of dusty side / non-dusty side and painted / non-painted for best adhesion. Surprisingly, putting the adhesive on the dusty side worked best. The material still does shed some dust in the wind, but it is reasonably rugged, more so than the fastblock insulation we were previously using, and it is only about $5 / square foot, which is a tenth the cost of the fastblock. It also sheds water fairly well. I was a bit surprised at how much of a difference insulating the tank made on our boiloff rates. I had been presuming that much of the boiloff was due to heat conduction from the rest of the 90 pound tanks that don’t get cooled that well during filling, but it turns out that the entire tanks get cooled a lot better with insulation. We used to have 10 psi in the tanks after filling, even with the vents open, but now it is just 3 psi and soon drops to 1 psi. We also insulated our test stand tank. http://media.armadilloaerospace.com/2006_08_12/insulating.jpg f94707b99f10d3d2ff44b033b507f82ddc6b860a 391 390 2012-11-09T15:27:43Z Vincent 1 /* Resources */ xcor and styrofoam wikitext text/x-wiki =Rocket Fuel tanks= [[Image:Restone_tank.jpg|right|320px]] In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. The sphere is the most lightweight volume (volume / area is minimized), but we can't have rockets as large as they are long, so cylinders with hemispheric caps are used. Having a single tank cut in two like for the Restone is efficient mass-wise and volume-wise but can bring new issues for insulation in case of a single cryogenic fluid (funny enough, that was the case for the Restone which used LOX and ethanol). The thicker insulation may overtake the mass benefits of a single tank. ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Wall thickness and material== Tank material first has to be stable with what's inside. Lists are available for cryogenic liquids at least. Besides this basic filter, the material choice mostly depends on money and on what's available on the market. For pressurized tanks, we will use 6061 aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. For example, aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. For a 0.2m wide tank, the thickness can be 1mm for the same pressure, or twice the pressure for the same thickness. ==Cryogenic fuel tanks== Cryogenic and also low boiling temperature liquids like nitrous oxide are persistently evaporating at ambient temperature. It's like having water at 100°C and providing always more heat to it. When the vapour pressure is high enough, and when tanks are solid enough too, the evaporation can reach [https://en.wikipedia.org/wiki/Evaporation#Evaporative_equilibrium equilibrium] and the tank can contain a stable mix of liquid and gas at high pressure. This is the case for nitrous oxide at temperatures below 36.4°C, its critical temperature above which it turns all into gas, no matter what pressure is used. The issue then becomes the density of the mixture, which drops greatly. For [[LOX]], the critical temperature is -118.59°C, and the critical pressure is 50.43 bar. There's no point in keeping it so much pressurized because it could just boil off at this temperature. Since the phase change occurs at a constant temperature, we can as well choose a temperature and a pressure at which the LOX density is high enough, but that's a trade-off with the evaporation rate. Since the temperature difference between inside and outside the tank is greater, even more heat is transferred to the LOX, and more evaporation is created. Tank insulation is then required to avoid venting all the propellant before actually using it (balloon or aircraft launch can take some time to get to the launch altitude). ===Calculating evaporation rate=== A material has a [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K), representing its ability to conduct heat. An insulation layer has a [https://en.wikipedia.org/wiki/R-value_(insulation) thermal resistance] (R-value) and its opposite, the [https://en.wikipedia.org/wiki/U-value#U-value thermal transmittance] (U-value), indicating how much resistance to heat the material provides. For an insulation layer of thickness ''L'', ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The heat transfer by convection is then given by the formula: Φ = A × U × (T1 - T2), in Watt. The thermal transmittance of the tank material can be ignored for first approximations. A is the area where the heat exchange is made, the outer surface of the insulation layer. T1 is the convecting air temperature, T2 is external tank temperature, which is close enough from the internal temperature (constant, boiling point given by pressure) to be taken for it. Examples can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc]. Finally, the evaporation rate is the heat of vaporization ΔH_vap/Φ in kg/s. That requires to know ΔH_vap for the chosen storage temperature, but graphs are available for common molecules like O2. ===Thermal insulation materials=== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of thermal conductivities is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. PU foam is a simple solution, it can be sprayed at the desired thickness but may be hard to spray in very thin layers. Aerogel is the best existing insulation material and it has been used by Armadillo Aerospace (see [[Rocket_Main_Tank#Resources|below]]). Aerogel for cryogenic applications is [http://www.aerogel.com/products/overview-product.html currently available] in 5mm or 10mm thick sheets. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |Aerogel |15 |13 |readily available in sheets, cheap |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayed |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | |readily available, cheap |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. ==Propellant lines== Pumps and Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. ==Resources== Armadillo Aerospace has [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=333 documented] their LOX tank insulation: We settled on using Pyrogel insulating blankets from Aspen Aerogels to insulate our tanks: [http://www.aerogel.com/products/pdf/Pyrogel_6250_DS.pdf#search=%22pyrogel%20insulation%22] After Phil figured out the right pattern to make the gores for the spherical tanks, it went very well. We used a spray adhesive to attach it, and we tested all the combinations of dusty side / non-dusty side and painted / non-painted for best adhesion. Surprisingly, putting the adhesive on the dusty side worked best. The material still does shed some dust in the wind, but it is reasonably rugged, more so than the fastblock insulation we were previously using, and it is only about $5 / square foot, which is a tenth the cost of the fastblock. It also sheds water fairly well. I was a bit surprised at how much of a difference insulating the tank made on our boiloff rates. I had been presuming that much of the boiloff was due to heat conduction from the rest of the 90 pound tanks that don’t get cooled that well during filling, but it turns out that the entire tanks get cooled a lot better with insulation. We used to have 10 psi in the tanks after filling, even with the vents open, but now it is just 3 psi and soon drops to 1 psi. We also insulated our test stand tank. http://media.armadilloaerospace.com/2006_08_12/insulating.jpg XCOR uses Styrofoam, a type of polystyrene with a 33 mW/m.K thermal conductivity, to insulate their aluminium LOX tanks. 65a32b24e9b0c6bf82010ca4220cdea3244eabaa 394 391 2012-11-09T15:57:45Z Vincent 1 /* Calculating evaporation rate */ introducing atmospheric model wikitext text/x-wiki =Rocket Fuel tanks= [[Image:Restone_tank.jpg|right|320px]] In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. The sphere is the most lightweight volume (volume / area is minimized), but we can't have rockets as large as they are long, so cylinders with hemispheric caps are used. Having a single tank cut in two like for the Restone is efficient mass-wise and volume-wise but can bring new issues for insulation in case of a single cryogenic fluid (funny enough, that was the case for the Restone which used LOX and ethanol). The thicker insulation may overtake the mass benefits of a single tank. ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Wall thickness and material== Tank material first has to be stable with what's inside. Lists are available for cryogenic liquids at least. Besides this basic filter, the material choice mostly depends on money and on what's available on the market. For pressurized tanks, we will use 6061 aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. For example, aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. For a 0.2m wide tank, the thickness can be 1mm for the same pressure, or twice the pressure for the same thickness. ==Cryogenic fuel tanks== Cryogenic and also low boiling temperature liquids like nitrous oxide are persistently evaporating at ambient temperature. It's like having water at 100°C and providing always more heat to it. When the vapour pressure is high enough, and when tanks are solid enough too, the evaporation can reach [https://en.wikipedia.org/wiki/Evaporation#Evaporative_equilibrium equilibrium] and the tank can contain a stable mix of liquid and gas at high pressure. This is the case for nitrous oxide at temperatures below 36.4°C, its critical temperature above which it turns all into gas, no matter what pressure is used. The issue then becomes the density of the mixture, which drops greatly. For [[LOX]], the critical temperature is -118.59°C, and the critical pressure is 50.43 bar. There's no point in keeping it so much pressurized because it could just boil off at this temperature. Since the phase change occurs at a constant temperature, we can as well choose a temperature and a pressure at which the LOX density is high enough, but that's a trade-off with the evaporation rate. Since the temperature difference between inside and outside the tank is greater, even more heat is transferred to the LOX, and more evaporation is created. Tank insulation is then required to avoid venting all the propellant before actually using it (balloon or aircraft launch can take some time to get to the launch altitude). ===Calculating evaporation rate=== A material has a [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K), representing its ability to conduct heat. An insulation layer has a [https://en.wikipedia.org/wiki/R-value_(insulation) thermal resistance] (R-value) and its opposite, the [https://en.wikipedia.org/wiki/U-value#U-value thermal transmittance] (U-value), indicating how much resistance to heat the material provides. For an insulation layer of thickness ''L'', ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The heat transfer by convection is then given by the formula: Φ = A × U × (T1 - T2), in Watt. The thermal transmittance of the tank material can be ignored for first approximations. A is the area where the heat exchange is made, the outer surface of the insulation layer. T1 is the convecting air temperature, T2 is external tank temperature, which is close enough from the internal temperature (constant, boiling point given by pressure) to be taken for it. Examples can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc]. Finally, the evaporation rate is the heat of vaporization ΔH_vap/Φ in kg/s. That requires to know ΔH_vap for the chosen storage temperature, but graphs are available for common molecules like O2. In the case of balloon or aircraft launches, the rocket stays filled with cryogenic fluids for a quite long time before being fired, with no possibility for refuelling. During the ascent to the launch altitude, the vehicle passes through different air layers with different temperatures and densities. A simple atmospheric model has to be made and used in order to calculate the heat transferred to the vehicle during the ascent, which is dominant over the heat transferred during powered flight. The ascent rate must also be approximated, possibly taken constant between some large levels. ===Thermal insulation materials=== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of thermal conductivities is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. PU foam is a simple solution, it can be sprayed at the desired thickness but may be hard to spray in very thin layers. Aerogel is the best existing insulation material and it has been used by Armadillo Aerospace (see [[Rocket_Main_Tank#Resources|below]]). Aerogel for cryogenic applications is [http://www.aerogel.com/products/overview-product.html currently available] in 5mm or 10mm thick sheets. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |Aerogel |15 |13 |readily available in sheets, cheap |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayed |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | |readily available, cheap |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. ==Propellant lines== Pumps and Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. ==Resources== Armadillo Aerospace has [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=333 documented] their LOX tank insulation: We settled on using Pyrogel insulating blankets from Aspen Aerogels to insulate our tanks: [http://www.aerogel.com/products/pdf/Pyrogel_6250_DS.pdf#search=%22pyrogel%20insulation%22] After Phil figured out the right pattern to make the gores for the spherical tanks, it went very well. We used a spray adhesive to attach it, and we tested all the combinations of dusty side / non-dusty side and painted / non-painted for best adhesion. Surprisingly, putting the adhesive on the dusty side worked best. The material still does shed some dust in the wind, but it is reasonably rugged, more so than the fastblock insulation we were previously using, and it is only about $5 / square foot, which is a tenth the cost of the fastblock. It also sheds water fairly well. I was a bit surprised at how much of a difference insulating the tank made on our boiloff rates. I had been presuming that much of the boiloff was due to heat conduction from the rest of the 90 pound tanks that don’t get cooled that well during filling, but it turns out that the entire tanks get cooled a lot better with insulation. We used to have 10 psi in the tanks after filling, even with the vents open, but now it is just 3 psi and soon drops to 1 psi. We also insulated our test stand tank. http://media.armadilloaerospace.com/2006_08_12/insulating.jpg XCOR uses Styrofoam, a type of polystyrene with a 33 mW/m.K thermal conductivity, to insulate their aluminium LOX tanks. 83cc8014a924d53265bcf11a019716c2a0e27963 401 394 2012-11-17T03:14:04Z Vincent 1 /* Calculating evaporation rate */ links to ISA wikitext text/x-wiki =Rocket Fuel tanks= [[Image:Restone_tank.jpg|right|320px]] In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. The sphere is the most lightweight volume (volume / area is minimized), but we can't have rockets as large as they are long, so cylinders with hemispheric caps are used. Having a single tank cut in two like for the Restone is efficient mass-wise and volume-wise but can bring new issues for insulation in case of a single cryogenic fluid (funny enough, that was the case for the Restone which used LOX and ethanol). The thicker insulation may overtake the mass benefits of a single tank. ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Wall thickness and material== Tank material first has to be stable with what's inside. Lists are available for cryogenic liquids at least. Besides this basic filter, the material choice mostly depends on money and on what's available on the market. For pressurized tanks, we will use 6061 aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. For example, aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. For a 0.2m wide tank, the thickness can be 1mm for the same pressure, or twice the pressure for the same thickness. ==Cryogenic fuel tanks== Cryogenic and also low boiling temperature liquids like nitrous oxide are persistently evaporating at ambient temperature. It's like having water at 100°C and providing always more heat to it. When the vapour pressure is high enough, and when tanks are solid enough too, the evaporation can reach [https://en.wikipedia.org/wiki/Evaporation#Evaporative_equilibrium equilibrium] and the tank can contain a stable mix of liquid and gas at high pressure. This is the case for nitrous oxide at temperatures below 36.4°C, its critical temperature above which it turns all into gas, no matter what pressure is used. The issue then becomes the density of the mixture, which drops greatly. For [[LOX]], the critical temperature is -118.59°C, and the critical pressure is 50.43 bar. There's no point in keeping it so much pressurized because it could just boil off at this temperature. Since the phase change occurs at a constant temperature, we can as well choose a temperature and a pressure at which the LOX density is high enough, but that's a trade-off with the evaporation rate. Since the temperature difference between inside and outside the tank is greater, even more heat is transferred to the LOX, and more evaporation is created. Tank insulation is then required to avoid venting all the propellant before actually using it (balloon or aircraft launch can take some time to get to the launch altitude). ===Calculating evaporation rate=== ''This section is under heavy work and information is not completely exact.'' A material has a [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K), representing its ability to conduct heat. An insulation layer has a [https://en.wikipedia.org/wiki/R-value_(insulation) thermal resistance] (R-value) and its opposite, the [https://en.wikipedia.org/wiki/U-value#U-value thermal transmittance] (U-value), indicating how much resistance to heat the material provides. For an insulation layer of thickness ''L'', ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The heat transfer by convection is then given by the formula: Φ = A × U × (T1 - T2), in Watt. The thermal transmittance of the tank material can be ignored for first approximations. A is the area where the heat exchange is made, the outer surface of the insulation layer. T1 is the convecting air temperature, T2 is external tank temperature, which is close enough from the internal temperature (constant, boiling point given by pressure) to be taken for it. Examples can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc]. Finally, the evaporation rate is the heat of vaporization ΔH_vap/Φ in kg/s. That requires to know ΔH_vap for the chosen storage temperature, but graphs are available for common molecules like O2. In the case of balloon or aircraft launches, the rocket stays filled with cryogenic fluids for a quite long time before being fired, with no possibility for refuelling. During the ascent to the launch altitude, the vehicle passes through different air layers with different temperatures and densities. A simple [[Flight_at_high_altitude#Gas_properties_and_altitude|atmospheric model]] has to be made (see [[File:ISA_atmospheric_model.c]]) and used in order to calculate the heat transferred to the vehicle during the ascent, which is dominant over the heat transferred during powered flight. The ascent rate will also be approximated, as constant between some gross altitude levels. ===Thermal insulation materials=== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of thermal conductivities is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. PU foam is a simple solution, it can be sprayed at the desired thickness but may be hard to spray in very thin layers. Aerogel is the best existing insulation material and it has been used by Armadillo Aerospace (see [[Rocket_Main_Tank#Resources|below]]). Aerogel for cryogenic applications is [http://www.aerogel.com/products/overview-product.html currently available] in 5mm or 10mm thick sheets. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |Aerogel |15 |13 |readily available in sheets, cheap |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayed |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | |readily available, cheap |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. ==Propellant lines== Pumps and Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. ==Resources== Armadillo Aerospace has [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=333 documented] their LOX tank insulation: We settled on using Pyrogel insulating blankets from Aspen Aerogels to insulate our tanks: [http://www.aerogel.com/products/pdf/Pyrogel_6250_DS.pdf#search=%22pyrogel%20insulation%22] After Phil figured out the right pattern to make the gores for the spherical tanks, it went very well. We used a spray adhesive to attach it, and we tested all the combinations of dusty side / non-dusty side and painted / non-painted for best adhesion. Surprisingly, putting the adhesive on the dusty side worked best. The material still does shed some dust in the wind, but it is reasonably rugged, more so than the fastblock insulation we were previously using, and it is only about $5 / square foot, which is a tenth the cost of the fastblock. It also sheds water fairly well. I was a bit surprised at how much of a difference insulating the tank made on our boiloff rates. I had been presuming that much of the boiloff was due to heat conduction from the rest of the 90 pound tanks that don’t get cooled that well during filling, but it turns out that the entire tanks get cooled a lot better with insulation. We used to have 10 psi in the tanks after filling, even with the vents open, but now it is just 3 psi and soon drops to 1 psi. We also insulated our test stand tank. http://media.armadilloaerospace.com/2006_08_12/insulating.jpg XCOR uses Styrofoam, a type of polystyrene with a 33 mW/m.K thermal conductivity, to insulate their aluminium LOX tanks. e5c9480219260bce9a48ef84f4ea85e3568b4d6b 6 72 385 2012-11-09T01:33:03Z Vincent 1 tank model for rocket mass versions 1 and 2 wikitext text/x-wiki tank model for rocket mass versions 1 and 2 e14a120531c9dc1569be39f9ffcfd1c14fe052dc 386 385 2012-11-09T01:36:36Z Vincent 1 uploaded a new version of "[[File:Rocket mass tanks v1-2.png]]":&#32;white bg wikitext text/x-wiki tank model for rocket mass versions 1 and 2 e14a120531c9dc1569be39f9ffcfd1c14fe052dc 6 73 387 2012-11-09T01:37:06Z Vincent 1 tank model for rocket mass version >= 3 wikitext text/x-wiki tank model for rocket mass version >= 3 da30cf08697fa4ee6420ee60dc013b52d0205054 0 55 396 340 2012-11-11T03:45:53Z Vincent 1 /* Approaches overview */ title changes and sr-71 link wikitext text/x-wiki =Flight at high altitude= Some information is summarized in the main page already, in the [[Main_Page#The_aircraft|aircraft section]]. ==Gas properties and altitude== [https://en.wikipedia.org/wiki/Density_of_air Density] is used to calculate [https://en.wikipedia.org/wiki/Lift_(force) lift] of an wing and [https://en.wikipedia.org/wiki/Thrust thrust] of an engine amongst other things. We absolutely need to know approximately what air densities will be faced in order to design everything. An atmosphere model should be used for design, but can also be used at runtime to verify that the actual atmosphere is within prediction range. Air density depends on pressure. [http://www.respirometry.org/look-up-table/barometric-pressure-vs-altitude This table] gives atmospheric pressure and temperature depending on altitude. We can see that a tenth of ground atmospheric pressure (ground-level: 1atm) is met at around 16km altitude, and a hundredth of it at around 31km altitude. Air density in the atmosphere is also related to the ratio of water vapour in it, as indicated on [https://wahiduddin.net/calc/density_altitude.htm this page]. The page also contains lots of formulas and calculators, most importantly the ''air density calculator'' that we'll use right below. Water vapour however, is much more rare when temperature goes down, as it does in the higher troposphere or low to mid stratosphere that we're aiming. The calculator gives us, with temperature and pressure values taken from the table mentioned above, values for density of: * 1.214 kg/m^3 at sea level (15°C) * 0.1877 kg/m^3 at 15km altitude (-57°C) * 0.0441 kg/m^3 at 25km altitude (-52°C) * 0.017 kg/m^3 at 30km altitude (-46°C) ==Turbofan engine's Mass flow rate calculation== One way of calculating the MFR is to use the [https://en.wikipedia.org/wiki/Continuity_equation#Fluid_dynamics continuity equation]. The mass of gas leaving the engine is the same than the mass of gas entering the engine, for which we know the density, plus the mass of the fuel, which is much lower than the mass of air. It's the velocity difference between input and output that creates the thrust. ==Approaches overview== ===Supersonic flight - high engine power=== '''Is it possible to have a low total pressure ratio engine operating at subsonic inlet speeds and low air density?''' The MiG 25 has supersonic inlet, which allows him to have a significant pressure increase before the compressor actually gives energy to the flow. A subsonic input air flow in the high-altitude conditions is likely to not provide enough oxygen for the combustion to maintain by itself, or a too poor mass flow rate to the turbine. The [https://en.wikipedia.org/wiki/Sr-71 SR-71] is another example for high-service ceiling (25900m, M3.2). ===Subsonic flight - high lift=== High engine power in low air density generally means supersonic flight, or at least, high flight speeds, which in return increase the lift of the aircraft or decrease its wingspan. Our next step is to make some calculations of the required winged area for subsonic low-density air travel, and assess the feasibility of our air launch to orbit project. 6a518806bb51b90c1795c4a2269edec24cb879d7 397 396 2012-11-11T03:49:31Z Vincent 1 /* Gas properties and altitude */ ISA wikitext text/x-wiki =Flight at high altitude= Some information is summarized in the main page already, in the [[Main_Page#The_aircraft|aircraft section]]. ==Gas properties and altitude== [https://en.wikipedia.org/wiki/Density_of_air Density] is used to calculate [https://en.wikipedia.org/wiki/Lift_(force) lift] of an wing and [https://en.wikipedia.org/wiki/Thrust thrust] of an engine amongst other things. We absolutely need to know approximately what air densities will be faced in order to design everything. An atmosphere model should be used for design, but can also be used at runtime to verify that the actual atmosphere is within prediction range. The most used model is the [https://en.wikipedia.org/wiki/International_Standard_Atmosphere ISA] (International Standard Atmosphere) from 1975. We [[Rocket_Main_Tank#Calculating_evaporation_rate|used]] this model in the heat flow modelling for cryogenic tanks. Air density depends on pressure. [http://www.respirometry.org/look-up-table/barometric-pressure-vs-altitude This table] gives atmospheric pressure and temperature depending on altitude. We can see that a tenth of ground atmospheric pressure (ground-level: 1atm) is met at around 16km altitude, and a hundredth of it at around 31km altitude. Air density in the atmosphere is also related to the ratio of water vapour in it, as indicated on [https://wahiduddin.net/calc/density_altitude.htm this page]. The page also contains lots of formulas and calculators, most importantly the ''air density calculator'' that we'll use right below. Water vapour however, is much more rare when temperature goes down, as it does in the higher troposphere or low to mid stratosphere that we're aiming. The calculator gives us, with temperature and pressure values taken from the table mentioned above, values for density of: * 1.214 kg/m^3 at sea level (15°C) * 0.1877 kg/m^3 at 15km altitude (-57°C) * 0.0441 kg/m^3 at 25km altitude (-52°C) * 0.017 kg/m^3 at 30km altitude (-46°C) ==Turbofan engine's Mass flow rate calculation== One way of calculating the MFR is to use the [https://en.wikipedia.org/wiki/Continuity_equation#Fluid_dynamics continuity equation]. The mass of gas leaving the engine is the same than the mass of gas entering the engine, for which we know the density, plus the mass of the fuel, which is much lower than the mass of air. It's the velocity difference between input and output that creates the thrust. ==Approaches overview== ===Supersonic flight - high engine power=== '''Is it possible to have a low total pressure ratio engine operating at subsonic inlet speeds and low air density?''' The MiG 25 has supersonic inlet, which allows him to have a significant pressure increase before the compressor actually gives energy to the flow. A subsonic input air flow in the high-altitude conditions is likely to not provide enough oxygen for the combustion to maintain by itself, or a too poor mass flow rate to the turbine. The [https://en.wikipedia.org/wiki/Sr-71 SR-71] is another example for high-service ceiling (25900m, M3.2). ===Subsonic flight - high lift=== High engine power in low air density generally means supersonic flight, or at least, high flight speeds, which in return increase the lift of the aircraft or decrease its wingspan. Our next step is to make some calculations of the required winged area for subsonic low-density air travel, and assess the feasibility of our air launch to orbit project. caeb5ccaaffe1e957fd313498a54214a4153f245 399 397 2012-11-17T02:46:39Z Vincent 1 /* Gas properties and altitude */ news about the atmospheric model wikitext text/x-wiki =Flight at high altitude= Some information is summarized in the main page already, in the [[Main_Page#The_aircraft|aircraft section]]. ==Gas properties and altitude== [https://en.wikipedia.org/wiki/Density_of_air Density] is used to calculate [https://en.wikipedia.org/wiki/Lift_(force) lift] of an wing and [https://en.wikipedia.org/wiki/Thrust thrust] of an engine amongst other things. We absolutely need to know approximately what air density will be faced in order to design everything. An atmosphere model should be used for design, but can also be used at runtime to verify that the actual atmosphere is within prediction range. The most used model is the [https://en.wikipedia.org/wiki/International_Standard_Atmosphere ISA] (International Standard Atmosphere) from 1975. It provides temperature, pressure and density as function of altitude, between -2km and 86km. We [[Rocket_Main_Tank#Calculating_evaporation_rate|used]] this model to estimate the heat transfer to cryogenic tanks during flight. Air density depends on pressure and thus on altitude. Some pressure values can be found in [http://www.respirometry.org/look-up-table/barometric-pressure-vs-altitude this table] or [http://www.engineeringtoolbox.com/international-standard-atmosphere-d_985.html this one]. We can see that a tenth of sea level atmospheric pressure (1atm: 101.325kPa) is met at around 16km altitude, and a hundredth of it at around 31km altitude. Air density in the atmosphere is also related to the ratio of water vapour in it, as indicated on [https://wahiduddin.net/calc/density_altitude.htm this page]. The page also contains lots of formulas and calculators, most importantly the ''air density calculator'' that we'll use right below for the examples. Water vapour is much less abundant when temperature goes down, as it does in the higher troposphere or low to mid stratosphere that we're aiming. The calculator gives us, with temperature and pressure values taken from the table mentioned above, values for density of: * 1.214 kg/m^3 at sea level (15°C) * 0.1877 kg/m^3 at 15km altitude (-57°C) * 0.0441 kg/m^3 at 25km altitude (-52°C) * 0.017 kg/m^3 at 30km altitude (-46°C) ==Turbofan engine's Mass flow rate calculation== One way of calculating the MFR is to use the [https://en.wikipedia.org/wiki/Continuity_equation#Fluid_dynamics continuity equation]. The mass of gas leaving the engine is the same than the mass of gas entering the engine, for which we know the density, plus the mass of the fuel, which is much lower than the mass of air. It's the velocity difference between input and output that creates the thrust. ==Approaches overview== ===Supersonic flight - high engine power=== '''Is it possible to have a low total pressure ratio engine operating at subsonic inlet speeds and low air density?''' The MiG 25 has supersonic inlet, which allows him to have a significant pressure increase before the compressor actually gives energy to the flow. A subsonic input air flow in the high-altitude conditions is likely to not provide enough oxygen for the combustion to maintain by itself, or a too poor mass flow rate to the turbine. The [https://en.wikipedia.org/wiki/Sr-71 SR-71] is another example for high-service ceiling (25900m, M3.2). ===Subsonic flight - high lift=== High engine power in low air density generally means supersonic flight, or at least, high flight speeds, which in return increase the lift of the aircraft or decrease its wingspan. Our next step is to make some calculations of the required winged area for subsonic low-density air travel, and assess the feasibility of our air launch to orbit project. 3d74e2fdc8cb6beccf74f19dc8f4413291b44e7b 6 74 400 2012-11-17T02:59:24Z Vincent 1 C Implementation of the international standard atmosphere ([https://en.wikipedia.org/wiki/International_Standard_Atmosphere ISA]) model. This is a port of http://www.digitaldutch.com/atmoscalc/ . wikitext text/x-wiki C Implementation of the international standard atmosphere ([https://en.wikipedia.org/wiki/International_Standard_Atmosphere ISA]) model. This is a port of http://www.digitaldutch.com/atmoscalc/ . 93e38ba5219a89f38ab32f12d0f7db609ad8c502 6 74 402 400 2012-11-17T03:17:37Z Vincent 1 comments on the source wikitext text/x-wiki C Implementation of the international standard atmosphere ([https://en.wikipedia.org/wiki/International_Standard_Atmosphere ISA]) model. This is a port of http://www.digitaldutch.com/atmoscalc/ . Unlike many sources found on the Internet, it supports the whole range of altitude defined by the ISA model, from -2km to 86km. A delta temperature is taken into account to adjust the sea level temperature if it's different from the model (15°C). 7f70ae0173e0a4bdaffc8a8f8ecdbc3bc0efc617 0 1 403 398 2012-11-19T00:39:26Z Vincent 1 /* News */ update for mid-nov wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''November 2012 update:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 150 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] have been created to represent the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. The program ([[File:Rocket_mass.c]]) modelling the mass of rockets is being refined to include cryogenic propellant evaporation. '''''August 2012 update:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may be weakly verified or partly false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. 159117dfad0cc2a28c313cff3d191193e11421e0 431 403 2012-11-25T02:58:35Z Vincent 1 /* News */ november update update wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''November 2012 update:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 150 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] have been created to illustrate the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. The program ([[File:Rocket_mass.c]]) modelling the mass of rockets is being refined and the cryogenic propellant vaporization has been evaluated by another program ([[File:Heat_transfer_to_rocket_tank.c]]). Both programs are freely available. '''''August 2012 update:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may be weakly verified or partly false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. 163986baa0edf7581aa17595f70dd4b1e288d8d3 438 431 2012-11-29T02:12:27Z Vincent 1 /* How to escape from Earth? */ gravityloss link to air breathers wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''November 2012 update:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 150 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] have been created to illustrate the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. The program ([[File:Rocket_mass.c]]) modelling the mass of rockets is being refined and the cryogenic propellant vaporization has been evaluated by another program ([[File:Heat_transfer_to_rocket_tank.c]]). Both programs are freely available. '''''August 2012 update:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may be weakly verified or partly false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. See [http://gravityloss.wordpress.com/2008/04/21/air-breathers-advantage this article] for more information on the differences of rocket and aircraft propulsion efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. 7186e8b13c1eafcdb3ea18ef91891ff30ef05489 440 438 2012-12-02T01:16:07Z Vincent 1 /* News */ updates for the erroneous gravity drag and thrust wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''December 2012 update:''''' A simple rocket trajectory model has been made in order to evaluate the trajectory of an aircraft- or balloon-released-rocket. It appears that the Delta V taken for granted for gravity drag for these rockets, around 800 m/s, is quite erroneous, or not possible with the expected overall thrust-to-weight ratio. Without taking into account the aerodynamic effects like lift and drag, for a ratio of around 1.5, the Delta V for gravity drag is at least 1300 m/s. To reduce it, the ratio should be higher, like 3.0, in that case it may be possible to have only 800 m/s Delta V for gravity, but the mass of the engine would be much higher, and the rocket is estimated to be at least 150kg more. The program evaluating the mass of the rocket has been refined, and adding the forgotten pipe linking the upper tank to the engine add approximately 100 kg of wet mass too... Incoming works will focus on how much we can expect the aerodynamic lift to compensate the gravity drag, and how much we can reduce the engine's thrust to decrease the mass of the vehicle while staying air-SSTO. '''''November 2012 update:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 150 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] have been created to illustrate the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. The program ([[File:Rocket_mass.c]]) modelling the mass of rockets is being refined and the cryogenic propellant vaporization has been evaluated by another program ([[File:Heat_transfer_to_rocket_tank.c]]). Both programs are freely available. '''''August 2012 update:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may be weakly verified or partly false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. See [http://gravityloss.wordpress.com/2008/04/21/air-breathers-advantage this article] for more information on the differences of rocket and aircraft propulsion efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. 9e20db6fd97bec5baab2133aa3602b3f09e5a196 0 55 404 399 2012-11-19T04:03:36Z Vincent 1 /* Gas properties and altitude */ more links for ISA wikitext text/x-wiki =Flight at high altitude= Some information is summarized in the main page already, in the [[Main_Page#The_aircraft|aircraft section]]. ==Gas properties and altitude== [https://en.wikipedia.org/wiki/Density_of_air Density] is a very useful property in aeronautics. It's for example used to calculate the [https://en.wikipedia.org/wiki/Lift_(force) lift] and the [https://en.wikipedia.org/wiki/Drag_(physics) drag] of an aerofoil, the [https://en.wikipedia.org/wiki/Thrust thrust] of an engine or even the heat transferred by air. We must know, at least approximately, what air density will be faced in order to design everything. An atmosphere model should be used for design, but can also be used at runtime to verify that the actual atmosphere is within prediction range. The most used model is the [https://en.wikipedia.org/wiki/International_Standard_Atmosphere ISA] (International Standard Atmosphere) from 1975. It is quite simple and provides temperature, pressure and density as function of altitude, between -2km and 86km. A calculator for the model is available [http://www-mdp.eng.cam.ac.uk/web/library/enginfo/aerothermal_dvd_only/aero/atmos/stdatm.html here] ([http://www-mdp.eng.cam.ac.uk/web/library/enginfo/aerothermal_dvd_only/aero/atmos/atmtab.html example values]). We [[Rocket_Main_Tank#Calculating_evaporation_rate|use]] this model to estimate the heat transfer to cryogenic tanks during rocket ascension to ignition altitude. Our implementation can be found here: [[File:ISA_atmospheric_model.c]]. Air density depends on pressure and temperature, and thus on altitude. Some pressure values can be found in [http://www.respirometry.org/look-up-table/barometric-pressure-vs-altitude this table] or [http://www.engineeringtoolbox.com/international-standard-atmosphere-d_985.html this one]. We can see that a tenth of sea level atmospheric pressure (1atm: 101.325kPa) is met at around 16km altitude, and a hundredth of it at around 31km altitude. Air density in the atmosphere is also related to the ratio of water vapour in it, as indicated on [https://wahiduddin.net/calc/density_altitude.htm this page]. The page also contains lots of formulas and calculators, most importantly the ''air density calculator'' that we'll use right below for the examples. Water vapour is much less abundant when temperature goes down, as it does in the higher troposphere or low to mid stratosphere that we're aiming. The calculator gives us, with temperature and pressure values taken from the table mentioned above, values for density of: * 1.214 kg/m^3 at sea level (15°C) * 0.1877 kg/m^3 at 15km altitude (-57°C) * 0.0441 kg/m^3 at 25km altitude (-52°C) * 0.017 kg/m^3 at 30km altitude (-46°C) ==Turbofan engine's Mass flow rate calculation== One way of calculating the MFR is to use the [https://en.wikipedia.org/wiki/Continuity_equation#Fluid_dynamics continuity equation]. The mass of gas leaving the engine is the same than the mass of gas entering the engine, for which we know the density, plus the mass of the fuel, which is much lower than the mass of air. It's the velocity difference between input and output that creates the thrust. ==Approaches overview== ===Supersonic flight - high engine power=== '''Is it possible to have a low total pressure ratio engine operating at subsonic inlet speeds and low air density?''' The MiG 25 has supersonic inlet, which allows him to have a significant pressure increase before the compressor actually gives energy to the flow. A subsonic input air flow in the high-altitude conditions is likely to not provide enough oxygen for the combustion to maintain by itself, or a too poor mass flow rate to the turbine. The [https://en.wikipedia.org/wiki/Sr-71 SR-71] is another example for high-service ceiling (25900m, M3.2). ===Subsonic flight - high lift=== High engine power in low air density generally means supersonic flight, or at least, high flight speeds, which in return increase the lift of the aircraft or decrease its wingspan. Our next step is to make some calculations of the required winged area for subsonic low-density air travel, and assess the feasibility of our air launch to orbit project. 7bc2aedd354d2e94b08d876e8d61f27b343d680d 0 24 405 401 2012-11-20T20:15:33Z Vincent 1 /* Calculating evaporation rate */ corrections and introductin of the new heat page wikitext text/x-wiki =Rocket Fuel tanks= [[Image:Restone_tank.jpg|right|320px]] In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. The sphere is the most lightweight volume (volume / area is minimized), but we can't have rockets as large as they are long, so cylinders with hemispheric caps are used. Having a single tank cut in two like for the Restone is efficient mass-wise and volume-wise but can bring new issues for insulation in case of a single cryogenic fluid (funny enough, that was the case for the Restone which used LOX and ethanol). The thicker insulation may overtake the mass benefits of a single tank. ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Wall thickness and material== Tank material first has to be stable with what's inside. Lists are available for cryogenic liquids at least. Besides this basic filter, the material choice mostly depends on money and on what's available on the market. For pressurized tanks, we will use 6061 aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. For example, aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. For a 0.2m wide tank, the thickness can be 1mm for the same pressure, or twice the pressure for the same thickness. ==Cryogenic fuel tanks== Cryogenic and also low boiling temperature liquids like nitrous oxide are persistently evaporating at ambient temperature. It's like having water at 100°C and providing always more heat to it. When the vapour pressure is high enough, and when tanks are solid enough too, the evaporation can reach [https://en.wikipedia.org/wiki/Evaporation#Evaporative_equilibrium equilibrium] and the tank can contain a stable mix of liquid and gas at high pressure. This is the case for nitrous oxide at temperatures below 36.4°C, its critical temperature above which it turns all into gas, no matter what pressure is used. The issue then becomes the density of the mixture, which drops greatly. For [[LOX]], the critical temperature is -118.59°C, and the critical pressure is 50.43 bar. There's no point in keeping it so much pressurized because it could just boil off at this temperature. Since the phase change occurs at a constant temperature, we can as well choose a temperature and a pressure at which the LOX density is high enough, but that's a trade-off with the evaporation rate. Since the temperature difference between inside and outside the tank is greater, even more heat is transferred to the LOX, and more evaporation is created. Tank insulation is then required to avoid venting all the propellant before actually using it (balloon or aircraft launch can take some time to get to the launch altitude). ===Calculating evaporation rate=== ''Main page: [[Heat transfer]]'' A material has a [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K), representing its ability to conduct heat. An insulation layer has a [https://en.wikipedia.org/wiki/R-value_(insulation) thermal resistance] (R-value) and its opposite, the [https://en.wikipedia.org/wiki/U-value#U-value thermal transmittance] (U-value), indicating how much resistance to heat the material provides in a particular use case. For an insulation layer of thickness ''L'', ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The heat transfer by convection is proportional to the temperature difference between the two parts, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], but do not include the calculations required for the convection heat transfer coefficient ''h'' that has to be known for air. Finally, the evaporation rate is the heat of vaporization ΔH_vap/Φ in kg/s. That requires to know ΔH_vap for the chosen storage temperature, but graphs are available for common molecules like O₂. In the case of balloon or aircraft launches, the rocket stays filled with cryogenic fluids for a quite long time before being fired, with no possibility for refuelling. During the ascent to the launch altitude, the vehicle passes through different air layers with different temperatures and densities. A simple [[Flight_at_high_altitude#Gas_properties_and_altitude|atmospheric model]] has to be made (see [[File:ISA_atmospheric_model.c]]) and used in order to calculate the heat transferred to the vehicle during the ascent, which is dominant over the heat transferred during powered flight. The ascent rate will also be approximated, as constant between some gross altitude levels. ===Thermal insulation materials=== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of thermal conductivities is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. PU foam is a simple solution, it can be sprayed at the desired thickness but may be hard to spray in very thin layers. Aerogel is the best existing insulation material and it has been used by Armadillo Aerospace (see [[Rocket_Main_Tank#Resources|below]]). Aerogel for cryogenic applications is [http://www.aerogel.com/products/overview-product.html currently available] in 5mm or 10mm thick sheets. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |Aerogel |15 |13 |readily available in sheets, cheap |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayed |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | |readily available, cheap |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. ==Propellant lines== Pumps and Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. ==Resources== Armadillo Aerospace has [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=333 documented] their LOX tank insulation: We settled on using Pyrogel insulating blankets from Aspen Aerogels to insulate our tanks: [http://www.aerogel.com/products/pdf/Pyrogel_6250_DS.pdf#search=%22pyrogel%20insulation%22] After Phil figured out the right pattern to make the gores for the spherical tanks, it went very well. We used a spray adhesive to attach it, and we tested all the combinations of dusty side / non-dusty side and painted / non-painted for best adhesion. Surprisingly, putting the adhesive on the dusty side worked best. The material still does shed some dust in the wind, but it is reasonably rugged, more so than the fastblock insulation we were previously using, and it is only about $5 / square foot, which is a tenth the cost of the fastblock. It also sheds water fairly well. I was a bit surprised at how much of a difference insulating the tank made on our boiloff rates. I had been presuming that much of the boiloff was due to heat conduction from the rest of the 90 pound tanks that don’t get cooled that well during filling, but it turns out that the entire tanks get cooled a lot better with insulation. We used to have 10 psi in the tanks after filling, even with the vents open, but now it is just 3 psi and soon drops to 1 psi. We also insulated our test stand tank. http://media.armadilloaerospace.com/2006_08_12/insulating.jpg XCOR uses Styrofoam, a type of polystyrene with a 33 mW/m.K thermal conductivity, to insulate their aluminium LOX tanks. 6ff2d961e8a2f980c0832d2092b85b394579f470 410 405 2012-11-21T00:22:10Z Vincent 1 /* Calculating evaporation rate */ update following the creation of the dedicated page on heat wikitext text/x-wiki =Rocket Fuel tanks= [[Image:Restone_tank.jpg|right|320px]] In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. The sphere is the most lightweight volume (volume / area is minimized), but we can't have rockets as large as they are long, so cylinders with hemispheric caps are used. Having a single tank cut in two like for the Restone is efficient mass-wise and volume-wise but can bring new issues for insulation in case of a single cryogenic fluid (funny enough, that was the case for the Restone which used LOX and ethanol). The thicker insulation may overtake the mass benefits of a single tank. ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Wall thickness and material== Tank material first has to be stable with what's inside. Lists are available for cryogenic liquids at least. Besides this basic filter, the material choice mostly depends on money and on what's available on the market. For pressurized tanks, we will use 6061 aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. For example, aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. For a 0.2m wide tank, the thickness can be 1mm for the same pressure, or twice the pressure for the same thickness. ==Cryogenic fuel tanks== Cryogenic and also low boiling temperature liquids like nitrous oxide are persistently evaporating at ambient temperature. It's like having water at 100°C and providing always more heat to it. When the vapour pressure is high enough, and when tanks are solid enough too, the evaporation can reach [https://en.wikipedia.org/wiki/Evaporation#Evaporative_equilibrium equilibrium] and the tank can contain a stable mix of liquid and gas at high pressure. This is the case for nitrous oxide at temperatures below 36.4°C, its critical temperature above which it turns all into gas, no matter what pressure is used. The issue then becomes the density of the mixture, which drops greatly. For [[LOX]], the critical temperature is -118.59°C, and the critical pressure is 50.43 bar. There's no point in keeping it so much pressurized because it could just boil off at this temperature. Since the phase change occurs at a constant temperature, we can as well choose a temperature and a pressure at which the LOX density is high enough, but that's a trade-off with the evaporation rate. Since the temperature difference between inside and outside the tank is greater, even more heat is transferred to the LOX, and more evaporation is created. Tank insulation is then required to avoid venting all the propellant before actually using it (balloon or aircraft launch can take some time to get to the launch altitude). ===Calculating evaporation rate=== ''Main page: [[Heat transfer]]'' In the case of balloon or aircraft launches, the rocket stays filled with cryogenic fluids for a quite long time before being fired, with no possibility for refuelling. First on the ground after tank filling, during the final pre-flight verifications, then during the ascent to the rocket launch altitude. The vehicle passes through different air layers with different temperatures and densities. A simple [[Flight_at_high_altitude#Gas_properties_and_altitude|atmospheric model]] has been made (see [[File:ISA_atmospheric_model.c]]) and will be used in order to calculate the heat transferred to the vehicle during the ascent, which is dominant over the heat transferred during powered flight. The ascent rate will also be approximated, as constant between some gross altitude levels. The evaporation rate can be calculated easily once the amount of heat provided by air is known, using the evaporation rate is the heat of vaporization ΔH_vap, the insulation layer's U-value, temperatures and so on. The issue is that the amount of heat provided by air is very hard to calculate. It should be done using CFD, but can be approximated in some conditions. The convection heat transfer coefficient ''h'' is the key concern. See the [[heat transfer]] page for more information. ===Thermal insulation materials=== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of thermal conductivities is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. PU foam is a simple solution, it can be sprayed at the desired thickness but may be hard to spray in very thin layers. Aerogel is the best existing insulation material and it has been used by Armadillo Aerospace (see [[Rocket_Main_Tank#Resources|below]]). Aerogel for cryogenic applications is [http://www.aerogel.com/products/overview-product.html currently available] in 5mm or 10mm thick sheets. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |Aerogel |15 |13 |readily available in sheets, cheap |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayed |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | |readily available, cheap |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. ==Propellant lines== Pumps and Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. ==Resources== Armadillo Aerospace has [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=333 documented] their LOX tank insulation: We settled on using Pyrogel insulating blankets from Aspen Aerogels to insulate our tanks: [http://www.aerogel.com/products/pdf/Pyrogel_6250_DS.pdf#search=%22pyrogel%20insulation%22] After Phil figured out the right pattern to make the gores for the spherical tanks, it went very well. We used a spray adhesive to attach it, and we tested all the combinations of dusty side / non-dusty side and painted / non-painted for best adhesion. Surprisingly, putting the adhesive on the dusty side worked best. The material still does shed some dust in the wind, but it is reasonably rugged, more so than the fastblock insulation we were previously using, and it is only about $5 / square foot, which is a tenth the cost of the fastblock. It also sheds water fairly well. I was a bit surprised at how much of a difference insulating the tank made on our boiloff rates. I had been presuming that much of the boiloff was due to heat conduction from the rest of the 90 pound tanks that don’t get cooled that well during filling, but it turns out that the entire tanks get cooled a lot better with insulation. We used to have 10 psi in the tanks after filling, even with the vents open, but now it is just 3 psi and soon drops to 1 psi. We also insulated our test stand tank. http://media.armadilloaerospace.com/2006_08_12/insulating.jpg XCOR uses Styrofoam, a type of polystyrene with a 33 mW/m.K thermal conductivity, to insulate their aluminium LOX tanks. 8fa6ef0b4e24269e288f7f352bddd0f71b0bf92a 425 410 2012-11-23T20:54:25Z Vincent 1 /* Calculating evaporation rate */ update following the heat calculations wikitext text/x-wiki =Rocket Fuel tanks= [[Image:Restone_tank.jpg|right|320px]] In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. The sphere is the most lightweight volume (volume / area is minimized), but we can't have rockets as large as they are long, so cylinders with hemispheric caps are used. Having a single tank cut in two like for the Restone is efficient mass-wise and volume-wise but can bring new issues for insulation in case of a single cryogenic fluid (funny enough, that was the case for the Restone which used LOX and ethanol). The thicker insulation may overtake the mass benefits of a single tank. ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Wall thickness and material== Tank material first has to be stable with what's inside. Lists are available for cryogenic liquids at least. Besides this basic filter, the material choice mostly depends on money and on what's available on the market. For pressurized tanks, we will use 6061 aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. For example, aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. For a 0.2m wide tank, the thickness can be 1mm for the same pressure, or twice the pressure for the same thickness. ==Cryogenic fuel tanks== Cryogenic and also low boiling temperature liquids like nitrous oxide are persistently evaporating at ambient temperature. It's like having water at 100°C and providing always more heat to it. When the vapour pressure is high enough, and when tanks are solid enough too, the evaporation can reach [https://en.wikipedia.org/wiki/Evaporation#Evaporative_equilibrium equilibrium] and the tank can contain a stable mix of liquid and gas at high pressure. This is the case for nitrous oxide at temperatures below 36.4°C, its critical temperature above which it turns all into gas, no matter what pressure is used. The issue then becomes the density of the mixture, which drops greatly. For [[LOX]], the critical temperature is -118.59°C, and the critical pressure is 50.43 bar. There's no point in keeping it so much pressurized because it could just boil off at this temperature. Since the phase change occurs at a constant temperature, we can as well choose a temperature and a pressure at which the LOX density is high enough, but that's a trade-off with the evaporation rate. Since the temperature difference between inside and outside the tank is greater, even more heat is transferred to the LOX, and more evaporation is created. Tank insulation is then required to avoid venting all the propellant before actually using it (balloon or aircraft launch can take some time to get to the launch altitude). ===Calculating evaporation rate=== ''Main page: [[Heat transfer]]'' In the case of balloon or aircraft launches, the rocket stays filled with cryogenic fluids for a quite long time before being fired, with no possibility for refuelling. First on the ground after tank filling, during the final pre-flight verifications, then during the ascent to the rocket ignition altitude. The vehicle passes through different air layers with different temperatures and densities. A simple [[Flight_at_high_altitude#Gas_properties_and_altitude|atmospheric model]] has been made (see [[File:ISA_atmospheric_model.c]]) and is used in order to calculate the heat transferred to the vehicle during the ascent, which is dominant over the heat transferred during the powered flight of the rocket. The ascent rate is approximated, as constant between some gross altitude levels. The evaporation rate can be calculated easily once the amount of heat provided by the surrounding air is known, using the heat of vaporization ΔH_vap of the propellant, the insulation layer's U-value, temperatures and so on. The issue is that the amount of heat provided by air is very hard to calculate. It should be done using CFD, but can be approximated in some conditions. The convection heat transfer coefficient ''h'' is the key concern. See the [[heat transfer]] page for more information and '''results''' for the LOX tank given by the first rocket mass [[Rocket:First_approximations#Our_first_approximation|approximation]]. ===Thermal insulation materials=== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of thermal conductivities is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. PU foam is a simple solution, it can be sprayed at the desired thickness but may be hard to spray in very thin layers. Aerogel is the best existing insulation material and it has been used by Armadillo Aerospace (see [[Rocket_Main_Tank#Resources|below]]). Aerogel for cryogenic applications is [http://www.aerogel.com/products/overview-product.html currently available] in 5mm or 10mm thick sheets. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |Aerogel |15 |13 |readily available in sheets, cheap |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayed |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | |readily available, cheap |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. ==Propellant lines== Pumps and Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. ==Resources== Armadillo Aerospace has [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=333 documented] their LOX tank insulation: We settled on using Pyrogel insulating blankets from Aspen Aerogels to insulate our tanks: [http://www.aerogel.com/products/pdf/Pyrogel_6250_DS.pdf#search=%22pyrogel%20insulation%22] After Phil figured out the right pattern to make the gores for the spherical tanks, it went very well. We used a spray adhesive to attach it, and we tested all the combinations of dusty side / non-dusty side and painted / non-painted for best adhesion. Surprisingly, putting the adhesive on the dusty side worked best. The material still does shed some dust in the wind, but it is reasonably rugged, more so than the fastblock insulation we were previously using, and it is only about $5 / square foot, which is a tenth the cost of the fastblock. It also sheds water fairly well. I was a bit surprised at how much of a difference insulating the tank made on our boiloff rates. I had been presuming that much of the boiloff was due to heat conduction from the rest of the 90 pound tanks that don’t get cooled that well during filling, but it turns out that the entire tanks get cooled a lot better with insulation. We used to have 10 psi in the tanks after filling, even with the vents open, but now it is just 3 psi and soon drops to 1 psi. We also insulated our test stand tank. http://media.armadilloaerospace.com/2006_08_12/insulating.jpg XCOR uses Styrofoam, a type of polystyrene with a 33 mW/m.K thermal conductivity, to insulate their aluminium LOX tanks. ea0c9f94c4f4337b31b92792c857751faf8772ff 0 75 406 2012-11-20T21:08:14Z Vincent 1 starting the page wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection== Heat convection occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. ==References== 5e1edf150b24672b9839d5d2d4fe1458250a1fc6 407 406 2012-11-20T22:40:23Z Vincent 1 more on heat transfers wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===The Nusselt number ''Nu''=== ===The Reynolds number ''Re''=== ===The Prandtl number ''Pr''=== ==Cases of application== ===Natural convection for horizontal cryogenic tank=== ===Forced convection on aircraft fuselage during flight=== ==References== '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 087f06350fb359b2bd2239407da615194e05939b 411 407 2012-11-21T01:45:12Z Vincent 1 /* Introduction to heat convection and conduction */ up to Nusselt wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = ''Nu''. The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Reynolds number ''Re''=== ===The Prandtl number ''Pr''=== ==Cases of application== ===Natural convection for horizontal cryogenic tank=== ===Forced convection on aircraft fuselage during flight=== ==References== '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] acfd8da53bfe71a1141863b59fb00f99b4976fff 412 411 2012-11-21T02:10:41Z Vincent 1 /* Cases of application */ starting the section wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = ''Nu''. The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Reynolds number ''Re''=== ===The Prandtl number ''Pr''=== ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of the cryogenic propellants on the ground and during ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. ===Natural convection for horizontal cryogenic tank=== On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C temperature, the density (rho) is 1.18391, and at the mean film temperature of -78.07°C, the viscosity µ is 1.30015e-05 Pa.s, the specific heat Cp is 1007.68 J/kg.K and the Prusselt number is 0.874462. ===Forced convection on aircraft fuselage during flight=== ==References== '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 80a7960bd03b7fe6170a1ac7803cb1afbeafc97b 413 412 2012-11-21T02:45:47Z Vincent 1 /* References */ 2nd book wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = ''Nu''. The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Reynolds number ''Re''=== ===The Prandtl number ''Pr''=== ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of the cryogenic propellants on the ground and during ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. ===Natural convection for horizontal cryogenic tank=== On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C temperature, the density (rho) is 1.18391, and at the mean film temperature of -78.07°C, the viscosity µ is 1.30015e-05 Pa.s, the specific heat Cp is 1007.68 J/kg.K and the Prusselt number is 0.874462. ===Forced convection on aircraft fuselage during flight=== ==References== '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 1554c64b9754d05683df7070b4ced431c7c9bcec 414 413 2012-11-21T04:05:39Z Vincent 1 nearly up to current status wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = ''Nu''. The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of the cryogenic propellants on the ground and during ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. ===Natural convection for horizontal cryogenic tank=== On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C temperature, the density (rho) is 1.18391, and at the mean film temperature of -78.07°C, the viscosity µ is 1.30015e-05 Pa.s, the specific heat Cp is 1007.68 J/kg.K and the Prusselt number is 0.874462. ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2] "Heat transfer in turbulent flow over a flat plate" is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr drops from 0.87 at sea level to 0.013 at 30km altitude, and the Reynolds number drops from 4.36714e+07 to 668455. If Re is nearly in the allowed range, Pr is far from it. We can't apply the Colburn analogy for altitudes above 2.5km, where Pr gets lower than 0.7. ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] f9e2d479af3577f29a61f04ad118203de6f00d74 415 414 2012-11-22T01:10:01Z Vincent 1 /* Introduction to heat convection and conduction */ Rayleigh number wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = Nu The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ===The Rayleigh number ''Ra''=== The [https://en.wikipedia.org/wiki/Rayleigh_Number Rayleigh number] is an indicator for natural convection heat transfer. When it is below the critical value for a fluid, heat transfer is primarily in the form of conduction; when it exceeds the critical value, heat transfer is primarily in the form of convection. {{SERVER}}/images/formulas_mirror/rayleigh_number_neg.png α is the [https://en.wikipedia.org/wiki/Thermal_diffusivity thermal diffusivity] {{SERVER}}/images/formulas_mirror/thermal_diffusivity_neg.png ; β is the [https://en.wikipedia.org/wiki/Coefficient_of_thermal_expansion thermal expansion coefficient], for an isobaric process it can be approximated to β = 1/T. ''x'' is the distance from the leading edge. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of the cryogenic propellants on the ground and during ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. ===Natural convection for horizontal cryogenic tank=== On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C temperature, the density (rho) is 1.18391, and at the mean film temperature of -78.07°C, the viscosity µ is 1.30015e-05 Pa.s, the specific heat Cp is 1007.68 J/kg.K and the Prusselt number is 0.874462. ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2] "Heat transfer in turbulent flow over a flat plate" is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr drops from 0.87 at sea level to 0.013 at 30km altitude, and the Reynolds number drops from 4.36714e+07 to 668455. If Re is nearly in the allowed range, Pr is far from it. We can't apply the Colburn analogy for altitudes above 2.5km, where Pr gets lower than 0.7. ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 774ede9f546d94c51ce92ed300e89a0b13e37501 416 415 2012-11-22T02:21:19Z Vincent 1 /* Cases of application */ calculation of h for ground operation wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = Nu The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ===The Rayleigh number ''Ra''=== The [https://en.wikipedia.org/wiki/Rayleigh_Number Rayleigh number] is an indicator for natural convection heat transfer. When it is below the critical value for a fluid, heat transfer is primarily in the form of conduction; when it exceeds the critical value, heat transfer is primarily in the form of convection. {{SERVER}}/images/formulas_mirror/rayleigh_number_neg.png α is the [https://en.wikipedia.org/wiki/Thermal_diffusivity thermal diffusivity] {{SERVER}}/images/formulas_mirror/thermal_diffusivity_neg.png ; β is the [https://en.wikipedia.org/wiki/Coefficient_of_thermal_expansion thermal expansion coefficient], for an isobaric process it can be approximated to β = 1/T. ''x'' is the distance from the leading edge. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of the cryogenic propellants on the ground and during ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. ===Natural convection for horizontal cryogenic tank=== On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. {{SERVER}}/images/formulas_mirror/h_natural_conv_external_horiz_cyl_neg.png D is the diameter of the approximated cylinder, in our case we take 0.266m, given by a [[Rocket:First_approximations#Minimum_mass_evaluation|first evaluation]] of the rocket mass program. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C or 298.15 K ground temperature: * the density (rho) is 1.18391 kg/m^3 and at the mean film temperature of -78.07°C or 195.075 K, given the propellant and tank temperature of 92 K: * the viscosity µ is 1.30015e-05 Pa.s (calculated using [http://www.cfd-online.com/Wiki/Sutherland's_law Sutherland's law]) * the kinematic viscosity nu is µ / rho = 1.0982e-5 m^2/s * the specific heat Cp is 1007.68 J/kg.K * the Prusselt number is 0.874462 * the thermal conductivity ''k'' is 0.0177375 W/m.K * α is thus k/(rho.Cp) = 1.4868e-5 m^2/s * β is 1/195.075 = 5.126e-3 * Ra_D is ((9.80665 * 5.126e-3) / (1.0982e-5 * 1.4868e-5)) * (92 - 298.15) * 0.266^3 = 6.3467e10 * 0.266^3 = 1.1945e9. * finally, ''h'' is 0.0177375/0.266 * (0.6 + (0.387 * 1.1945e9^1/6) / (1 + (0.559/0.874462)^9/16)^8/27)^2 = 0.024382 W/m^2.K ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2], ''Heat transfer in turbulent flow over a flat plate'', is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr drops from 0.87 at sea level to 0.013 at 30km altitude, and the Reynolds number drops from 4.36714e+07 to 668455. If Re is nearly in the allowed range, Pr is far from it. We can't apply the Colburn analogy for altitudes above 2.5km, where Pr gets lower than 0.7. ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] e3f335ef85efdf9a8080068acd80ef667d5cc51d 417 416 2012-11-22T03:59:24Z Vincent 1 /* Natural convection for horizontal cryogenic tank */ first approximation and references wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = Nu The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ===The Rayleigh number ''Ra''=== The [https://en.wikipedia.org/wiki/Rayleigh_Number Rayleigh number] is an indicator for natural convection heat transfer. When it is below the critical value for a fluid, heat transfer is primarily in the form of conduction; when it exceeds the critical value, heat transfer is primarily in the form of convection. {{SERVER}}/images/formulas_mirror/rayleigh_number_neg.png α is the [https://en.wikipedia.org/wiki/Thermal_diffusivity thermal diffusivity] {{SERVER}}/images/formulas_mirror/thermal_diffusivity_neg.png ; β is the [https://en.wikipedia.org/wiki/Coefficient_of_thermal_expansion thermal expansion coefficient], for an isobaric process it can be approximated to β = 1/T. ''x'' is the distance from the leading edge. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of the cryogenic propellants on the ground and during ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. ===Natural convection for horizontal cryogenic tank=== '''''The results presented in this section are incorrect but the errors have not been identified yet.''''' On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. {{SERVER}}/images/formulas_mirror/h_natural_conv_external_horiz_cyl_neg.png D is the diameter of the approximated cylinder, in our case we take 0.266m, given by a [[Rocket:First_approximations#Minimum_mass_evaluation|first evaluation]] of the rocket mass program. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C or 298.15 K ground temperature: * the density (rho) is 1.18391 kg/m^3 and at the mean film temperature of -78.07°C or 195.075 K, given the propellant and tank temperature of 92 K: * the viscosity µ is 1.30015e-05 Pa.s (calculated using [http://www.cfd-online.com/Wiki/Sutherland's_law Sutherland's law]) * the kinematic viscosity nu is µ / rho = 1.0982e-5 m^2/s * the specific heat Cp is 1007.68 J/kg.K (approximated using [http://ninova.itu.edu.tr/tr/dersler/ucak-uzay-fakultesi/965/uck-421/ekkaynaklar?g96162 this interpolation]) * the Prandtl number is 0.874462 * the thermal conductivity of air ''k'' is 0.0177375 W/m.K (approximated using [http://physics.tutorvista.com/heat/heat-transfer.html#thermal-conductivity-of-air this interpolation]) * α is thus k/(rho.Cp) = 1.4868e-5 m^2/s * β is 1/195.075 = 5.126e-3 1/K * Ra_D is ((9.80665 * 5.126e-3) / (1.0982e-5 * 1.4868e-5)) * (298.15 - 92) * 0.266^3 = 6.3467e10 * 0.266^3 = 1.1945e9. * finally, ''h'' is 0.0177375/0.266 * (0.6 + (0.387 * 1.1945e9^1/6) / (1 + (0.559/0.874462)^9/16)^8/27)^2 '''= 0.024382 W/m^2.K''' We finally have a heat transfer coefficient for air. We can compute the total heat transferred from it, using Newton's law equation [[Heat_transfer#Heat_transfer_coefficient_h|(top)]]. We need to specify the area on which the heat transfer will apply, the approximated cylinder, and the number of seconds during which the heat is transferred. Tanks are a cylinder with two hemispherical end-caps. Since the end-caps still conduct heat, we will include their area and assume it is part of the approximed cylinder's area. Assumed area is then 1.362m^2. Let's take 600 seconds (10 minutes) for the time spent on the ground between tanks filling and aircraft lift-off. We assume the temperature to be constant on the ground and in the tank while heat is transferred. This is correct since the heat energy feeds a phase change in the propellant (vaporization) and not a temperature increase of some material. ''dQ/dt'' = ''h'' * ''A'' * (''T_ground'' - ''T_prop'') = 0.024382 * 1.362 * (298.15 - 92) = 6.846 J. We can now [[Rocket_Main_Tank#Calculating_evaporation_rate|calculate]] the evaporation rate of the propellant with and without insulation. The heat of vaporization for [[LOX]] is 213 kJ/kg. Without insulation, the rate of vaporization is 32mg/s, making 19 grammes of LOX evaporated in 600s. That does not seems quite right... With a 10mm [[Rocket_Main_Tank#Thermal_insulation_materials|Aerogel]] insulation, the thermal conductivity being 15 mW/m.K, its U-value being k/L = 0.015/0.010 = 1.5 W/m^2.K. The rate of vaporization is then? ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2], ''Heat transfer in turbulent flow over a flat plate'', is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr drops from 0.87 at sea level to 0.013 at 30km altitude, and the Reynolds number drops from 4.36714e+07 to 668455. If Re is nearly in the allowed range, Pr is far from it. We can't apply the Colburn analogy for altitudes above 2.5km, where Pr gets lower than 0.7. ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] b8721c3f0ad51d571d3cbb6f5f6f08a3a836824d 418 417 2012-11-23T01:36:40Z Vincent 1 /* References */ lecture of sezai wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = Nu The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ===The Rayleigh number ''Ra''=== The [https://en.wikipedia.org/wiki/Rayleigh_Number Rayleigh number] is an indicator for natural convection heat transfer. When it is below the critical value for a fluid, heat transfer is primarily in the form of conduction; when it exceeds the critical value, heat transfer is primarily in the form of convection. {{SERVER}}/images/formulas_mirror/rayleigh_number_neg.png α is the [https://en.wikipedia.org/wiki/Thermal_diffusivity thermal diffusivity] {{SERVER}}/images/formulas_mirror/thermal_diffusivity_neg.png ; β is the [https://en.wikipedia.org/wiki/Coefficient_of_thermal_expansion thermal expansion coefficient], for an isobaric process it can be approximated to β = 1/T. ''x'' is the distance from the leading edge. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of the cryogenic propellants on the ground and during ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. ===Natural convection for horizontal cryogenic tank=== '''''The results presented in this section are incorrect but the errors have not been identified yet.''''' On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. {{SERVER}}/images/formulas_mirror/h_natural_conv_external_horiz_cyl_neg.png D is the diameter of the approximated cylinder, in our case we take 0.266m, given by a [[Rocket:First_approximations#Minimum_mass_evaluation|first evaluation]] of the rocket mass program. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C or 298.15 K ground temperature: * the density (rho) is 1.18391 kg/m^3 and at the mean film temperature of -78.07°C or 195.075 K, given the propellant and tank temperature of 92 K: * the viscosity µ is 1.30015e-05 Pa.s (calculated using [http://www.cfd-online.com/Wiki/Sutherland's_law Sutherland's law]) * the kinematic viscosity nu is µ / rho = 1.0982e-5 m^2/s * the specific heat Cp is 1007.68 J/kg.K (approximated using [http://ninova.itu.edu.tr/tr/dersler/ucak-uzay-fakultesi/965/uck-421/ekkaynaklar?g96162 this interpolation]) * the Prandtl number is 0.874462 * the thermal conductivity of air ''k'' is 0.0177375 W/m.K (approximated using [http://physics.tutorvista.com/heat/heat-transfer.html#thermal-conductivity-of-air this interpolation]) * α is thus k/(rho.Cp) = 1.4868e-5 m^2/s * β is 1/195.075 = 5.126e-3 1/K * Ra_D is ((9.80665 * 5.126e-3) / (1.0982e-5 * 1.4868e-5)) * (298.15 - 92) * 0.266^3 = 6.3467e10 * 0.266^3 = 1.1945e9. * finally, ''h'' is 0.0177375/0.266 * (0.6 + (0.387 * 1.1945e9^1/6) / (1 + (0.559/0.874462)^9/16)^8/27)^2 '''= 0.024382 W/m^2.K''' We finally have a heat transfer coefficient for air. We can compute the total heat transferred from it, using Newton's law equation [[Heat_transfer#Heat_transfer_coefficient_h|(top)]]. We need to specify the area on which the heat transfer will apply, the approximated cylinder, and the number of seconds during which the heat is transferred. Tanks are a cylinder with two hemispherical end-caps. Since the end-caps still conduct heat, we will include their area and assume it is part of the approximed cylinder's area. Assumed area is then 1.362m^2. Let's take 600 seconds (10 minutes) for the time spent on the ground between tanks filling and aircraft lift-off. We assume the temperature to be constant on the ground and in the tank while heat is transferred. This is correct since the heat energy feeds a phase change in the propellant (vaporization) and not a temperature increase of some material. ''dQ/dt'' = ''h'' * ''A'' * (''T_ground'' - ''T_prop'') = 0.024382 * 1.362 * (298.15 - 92) = 6.846 J. We can now [[Rocket_Main_Tank#Calculating_evaporation_rate|calculate]] the evaporation rate of the propellant with and without insulation. The heat of vaporization for [[LOX]] is 213 kJ/kg. Without insulation, the rate of vaporization is 32mg/s, making 19 grammes of LOX evaporated in 600s. That does not seems quite right... With a 10mm [[Rocket_Main_Tank#Thermal_insulation_materials|Aerogel]] insulation, the thermal conductivity being 15 mW/m.K, its U-value being k/L = 0.015/0.010 = 1.5 W/m^2.K. The rate of vaporization is then? ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2], ''Heat transfer in turbulent flow over a flat plate'', is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr drops from 0.87 at sea level to 0.013 at 30km altitude, and the Reynolds number drops from 4.36714e+07 to 668455. If Re is nearly in the allowed range, Pr is far from it. We can't apply the Colburn analogy for altitudes above 2.5km, where Pr gets lower than 0.7. ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 3. Lecture on '''Advanced Heat Transfer''', Ibrahim Sezai, 2011. Distributed as open courseware, [http://opencourses.emu.edu.tr/course/view.php?id=23 here] e74de61d02ec0f7d514b10e3fe88d96bb167e538 419 418 2012-11-23T03:56:24Z Vincent 1 /* Forced convection on aircraft fuselage during flight */ found the bug int the code... Pr is fine. First approximation of the flight vaporization rate wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = Nu The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ===The Rayleigh number ''Ra''=== The [https://en.wikipedia.org/wiki/Rayleigh_Number Rayleigh number] is an indicator for natural convection heat transfer. When it is below the critical value for a fluid, heat transfer is primarily in the form of conduction; when it exceeds the critical value, heat transfer is primarily in the form of convection. {{SERVER}}/images/formulas_mirror/rayleigh_number_neg.png α is the [https://en.wikipedia.org/wiki/Thermal_diffusivity thermal diffusivity] {{SERVER}}/images/formulas_mirror/thermal_diffusivity_neg.png ; β is the [https://en.wikipedia.org/wiki/Coefficient_of_thermal_expansion thermal expansion coefficient], for an isobaric process it can be approximated to β = 1/T. ''x'' is the distance from the leading edge. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of the cryogenic propellants on the ground and during ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. ===Natural convection for horizontal cryogenic tank=== '''''The results presented in this section are incorrect but the errors have not been identified yet.''''' On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. {{SERVER}}/images/formulas_mirror/h_natural_conv_external_horiz_cyl_neg.png D is the diameter of the approximated cylinder, in our case we take 0.266m, given by a [[Rocket:First_approximations#Minimum_mass_evaluation|first evaluation]] of the rocket mass program. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C or 298.15 K ground temperature: * the density (rho) is 1.18391 kg/m^3 and at the mean film temperature of -78.07°C or 195.075 K, given the propellant and tank temperature of 92 K: * the viscosity µ is 1.30015e-05 Pa.s (calculated using [http://www.cfd-online.com/Wiki/Sutherland's_law Sutherland's law]) * the kinematic viscosity nu is µ / rho = 1.0982e-5 m^2/s * the specific heat Cp is 1007.68 J/kg.K (approximated using [http://ninova.itu.edu.tr/tr/dersler/ucak-uzay-fakultesi/965/uck-421/ekkaynaklar?g96162 this interpolation]) * the Prandtl number is 0.874462 * the thermal conductivity of air ''k'' is 0.0177375 W/m.K (approximated using [http://physics.tutorvista.com/heat/heat-transfer.html#thermal-conductivity-of-air this interpolation]) * α is thus k/(rho.Cp) = 1.4868e-5 m^2/s * β is 1/195.075 = 5.126e-3 1/K * Ra_D is ((9.80665 * 5.126e-3) / (1.0982e-5 * 1.4868e-5)) * (298.15 - 92) * 0.266^3 = 6.3467e10 * 0.266^3 = 1.1945e9. * finally, ''h'' is 0.0177375/0.266 * (0.6 + (0.387 * 1.1945e9^1/6) / (1 + (0.559/0.874462)^9/16)^8/27)^2 '''= 0.024382 W/m^2.K''' We finally have a heat transfer coefficient for air. We can compute the total heat transferred from it, using Newton's law equation [[Heat_transfer#Heat_transfer_coefficient_h|(top)]]. We need to specify the area on which the heat transfer will apply, the approximated cylinder, and the number of seconds during which the heat is transferred. Tanks are a cylinder with two hemispherical end-caps. Since the end-caps still conduct heat, we will include their area and assume it is part of the approximed cylinder's area. Assumed area is then 1.362m^2. Let's take 600 seconds (10 minutes) for the time spent on the ground between tanks filling and aircraft lift-off. We assume the temperature to be constant on the ground and in the tank while heat is transferred. This is correct since the heat energy feeds a phase change in the propellant (vaporization) and not a temperature increase of some material. ''dQ/dt'' = ''h'' * ''A'' * (''T_ground'' - ''T_prop'') = 0.024382 * 1.362 * (298.15 - 92) = 6.846 J. We can now [[Rocket_Main_Tank#Calculating_evaporation_rate|calculate]] the evaporation rate of the propellant with and without insulation. The heat of vaporization for [[LOX]] is 213 kJ/kg. Without insulation, the rate of vaporization is 32mg/s, making 19 grammes of LOX evaporated in 600s. That does not seems quite right... With a 10mm [[Rocket_Main_Tank#Thermal_insulation_materials|Aerogel]] insulation, the thermal conductivity being 15 mW/m.K, its U-value being k/L = 0.015/0.010 = 1.5 W/m^2.K. The rate of vaporization is then? ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2], ''Heat transfer in turbulent flow over a flat plate'', is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr varies between 0.73 and 0.75, and the Reynolds number for the characteristic length L of the tank (Re_L) drops from 4.36714e+07 to 6.68455e+06 with altitude. This value of Re confirms that the boundary layer is turbulent. [2] indicates that the Colburn analogy can still be used for Pr up to 10^8, but with some loss of accuracy. We use the following approximation to calculate the average Nusselt number for forced convection in turbulent flows: Nu_L = 0.0366 (Re_L^0.8 - 14500) Pr^1/3 given in [2]. The calculated average Nusselt number for ascent operation is 42353 at sea level and drops to 1037 at 30km altitude. ''h_L'' can then be calculated too, ''h_L'' = Nu_L * ''k''_air / L. It varies from 460.8 on the ground to 9.6 W/m^2.K at 30km altitude. The simple flight path approximation is used to calculate the average ''h'' value for every second of the flight. We also need to calculate the dQ/dt value for each second, since the external temperature changes as the aircraft ascends. The integrated value of heat Q is finally: Q = 1144.12 kJ. '''For [[LOX]], it translates into more than 55kg of oxidizer evaporated during the ascent, when no insulation is used.''' ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 3. Lecture on '''Advanced Heat Transfer''', Ibrahim Sezai, 2011. Distributed as open courseware, [http://opencourses.emu.edu.tr/course/view.php?id=23 here] 5a204d53d40d2f8bc92a2e1273e3f57ecdfc89e2 420 419 2012-11-23T04:08:55Z Vincent 1 /* Forced convection on aircraft fuselage during flight */ first approx of the vaporization rate with an insulation layer wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = Nu The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ===The Rayleigh number ''Ra''=== The [https://en.wikipedia.org/wiki/Rayleigh_Number Rayleigh number] is an indicator for natural convection heat transfer. When it is below the critical value for a fluid, heat transfer is primarily in the form of conduction; when it exceeds the critical value, heat transfer is primarily in the form of convection. {{SERVER}}/images/formulas_mirror/rayleigh_number_neg.png α is the [https://en.wikipedia.org/wiki/Thermal_diffusivity thermal diffusivity] {{SERVER}}/images/formulas_mirror/thermal_diffusivity_neg.png ; β is the [https://en.wikipedia.org/wiki/Coefficient_of_thermal_expansion thermal expansion coefficient], for an isobaric process it can be approximated to β = 1/T. ''x'' is the distance from the leading edge. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of the cryogenic propellants on the ground and during ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. ===Natural convection for horizontal cryogenic tank=== '''''The results presented in this section are incorrect but the errors have not been identified yet.''''' On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. {{SERVER}}/images/formulas_mirror/h_natural_conv_external_horiz_cyl_neg.png D is the diameter of the approximated cylinder, in our case we take 0.266m, given by a [[Rocket:First_approximations#Minimum_mass_evaluation|first evaluation]] of the rocket mass program. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C or 298.15 K ground temperature: * the density (rho) is 1.18391 kg/m^3 and at the mean film temperature of -78.07°C or 195.075 K, given the propellant and tank temperature of 92 K: * the viscosity µ is 1.30015e-05 Pa.s (calculated using [http://www.cfd-online.com/Wiki/Sutherland's_law Sutherland's law]) * the kinematic viscosity nu is µ / rho = 1.0982e-5 m^2/s * the specific heat Cp is 1007.68 J/kg.K (approximated using [http://ninova.itu.edu.tr/tr/dersler/ucak-uzay-fakultesi/965/uck-421/ekkaynaklar?g96162 this interpolation]) * the Prandtl number is 0.874462 * the thermal conductivity of air ''k'' is 0.0177375 W/m.K (approximated using [http://physics.tutorvista.com/heat/heat-transfer.html#thermal-conductivity-of-air this interpolation]) * α is thus k/(rho.Cp) = 1.4868e-5 m^2/s * β is 1/195.075 = 5.126e-3 1/K * Ra_D is ((9.80665 * 5.126e-3) / (1.0982e-5 * 1.4868e-5)) * (298.15 - 92) * 0.266^3 = 6.3467e10 * 0.266^3 = 1.1945e9. * finally, ''h'' is 0.0177375/0.266 * (0.6 + (0.387 * 1.1945e9^1/6) / (1 + (0.559/0.874462)^9/16)^8/27)^2 '''= 0.024382 W/m^2.K''' We finally have a heat transfer coefficient for air. We can compute the total heat transferred from it, using Newton's law equation [[Heat_transfer#Heat_transfer_coefficient_h|(top)]]. We need to specify the area on which the heat transfer will apply, the approximated cylinder, and the number of seconds during which the heat is transferred. Tanks are a cylinder with two hemispherical end-caps. Since the end-caps still conduct heat, we will include their area and assume it is part of the approximed cylinder's area. Assumed area is then 1.362m^2. Let's take 600 seconds (10 minutes) for the time spent on the ground between tanks filling and aircraft lift-off. We assume the temperature to be constant on the ground and in the tank while heat is transferred. This is correct since the heat energy feeds a phase change in the propellant (vaporization) and not a temperature increase of some material. ''dQ/dt'' = ''h'' * ''A'' * (''T_ground'' - ''T_prop'') = 0.024382 * 1.362 * (298.15 - 92) = 6.846 J. We can now [[Rocket_Main_Tank#Calculating_evaporation_rate|calculate]] the evaporation rate of the propellant with and without insulation. The heat of vaporization for [[LOX]] is 213 kJ/kg. Without insulation, the rate of vaporization is 32mg/s, making 19 grammes of LOX evaporated in 600s. That does not seems quite right... With a 10mm [[Rocket_Main_Tank#Thermal_insulation_materials|Aerogel]] insulation, the thermal conductivity being 15 mW/m.K, its U-value being k/L = 0.015/0.010 = 1.5 W/m^2.K. The rate of vaporization is then? ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2], ''Heat transfer in turbulent flow over a flat plate'', is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr varies between 0.73 and 0.75, and the Reynolds number for the characteristic length L of the tank (Re_L) drops from 4.36714e+07 to 6.68455e+06 with altitude. This value of Re confirms that the boundary layer is turbulent. [2] indicates that the Colburn analogy can still be used for Pr up to 10^8, but with some loss of accuracy. We use the following approximation to calculate the average Nusselt number for forced convection in turbulent flows: Nu_L = 0.0366 (Re_L^0.8 - 14500) Pr^1/3 given in [2]. The calculated average Nusselt number for ascent operation is 42353 at sea level and drops to 1037 at 30km altitude. ''h_L'' can then be calculated too, ''h_L'' = Nu_L * ''k''_air / L. It varies from 460.8 on the ground to 9.6 W/m^2.K at 30km altitude. The simple flight path approximation is used to calculate the average ''h'' value for every second of the flight. We also need to calculate the dQ/dt value for each second, since the external temperature changes as the aircraft ascends. The integrated value of heat Q is finally: Q = 1144.12 kJ. '''For [[LOX]], it translates into more than 53kg of oxidizer evaporated during the ascent, when no insulation is used. With an Aerogel insulation layer of 10mm, the mass of evaporated LOX is evaluated to 0.87kg.''' That corresponds to Q = 185.41 kJ. The overall heat transfer coefficient U was calculated using [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#Alternative_Method_.28A_simple_method_for_determining_the_overall_heat_transfer_coefficient.29 this] method. ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 3. Lecture on '''Advanced Heat Transfer''', Ibrahim Sezai, 2011. Distributed as open courseware, [http://opencourses.emu.edu.tr/course/view.php?id=23 here] 6690c7c99219286c90268bdd6eaa88f81e241ad3 421 420 2012-11-23T04:18:12Z Vincent 1 /* Natural convection for horizontal cryogenic tank */ fixing text for the h error wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = Nu The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ===The Rayleigh number ''Ra''=== The [https://en.wikipedia.org/wiki/Rayleigh_Number Rayleigh number] is an indicator for natural convection heat transfer. When it is below the critical value for a fluid, heat transfer is primarily in the form of conduction; when it exceeds the critical value, heat transfer is primarily in the form of convection. {{SERVER}}/images/formulas_mirror/rayleigh_number_neg.png α is the [https://en.wikipedia.org/wiki/Thermal_diffusivity thermal diffusivity] {{SERVER}}/images/formulas_mirror/thermal_diffusivity_neg.png ; β is the [https://en.wikipedia.org/wiki/Coefficient_of_thermal_expansion thermal expansion coefficient], for an isobaric process it can be approximated to β = 1/T. ''x'' is the distance from the leading edge. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of the cryogenic propellants on the ground and during ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. ===Natural convection for horizontal cryogenic tank=== '''''The results presented in this section are not yet correct. An error has been identified in the calculation of h, but it may not be the only one.''''' On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. {{SERVER}}/images/formulas_mirror/h_natural_conv_external_horiz_cyl_neg.png D is the diameter of the approximated cylinder, in our case we take 0.266m, given by a [[Rocket:First_approximations#Minimum_mass_evaluation|first evaluation]] of the rocket mass program. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C or 298.15 K ground temperature: * the density (rho) is 1.18391 kg/m^3 and at the mean film temperature of -78.07°C or 195.075 K, given the propellant and tank temperature of 92 K: * the viscosity µ is 1.30015e-05 Pa.s (calculated using [http://www.cfd-online.com/Wiki/Sutherland's_law Sutherland's law]) * the kinematic viscosity nu is µ / rho = 1.0982e-5 m^2/s * the specific heat Cp is 1007.68 J/kg.K (approximated using [http://ninova.itu.edu.tr/tr/dersler/ucak-uzay-fakultesi/965/uck-421/ekkaynaklar?g96162 this interpolation]) * the Prandtl number is 0.874462 * the thermal conductivity of air ''k'' is 0.0177375 W/m.K (approximated using [http://physics.tutorvista.com/heat/heat-transfer.html#thermal-conductivity-of-air this interpolation]) * α is thus k/(rho.Cp) = 1.4868e-5 m^2/s * β is approximated to 1/T = 1/195.075 = 5.126e-3 1/K * Ra_D is ((9.80665 * 5.126e-3) / (1.0982e-5 * 1.4868e-5)) * (298.15 - 92) * 0.266^3 = 6.3467e10 * 0.266^3 = 1.1945e9. * finally, ''h'' is 0.0177375/0.266 * (0.6 + (0.387 * 1.1945e9^1/6) / (1 + (0.559/0.874462)^9/16)^8/27)^2 '''= 0.024382 W/m^2.K''' . '''''WARNING: h should bee much higher.''''' We finally have a heat transfer coefficient for air. We can compute the total heat transferred from it, using Newton's law equation [[Heat_transfer#Heat_transfer_coefficient_h|(top)]]. We need to specify the area on which the heat transfer will apply, the approximated cylinder, and the number of seconds during which the heat is transferred. Tanks are a cylinder with two hemispherical end-caps. Since the end-caps still conduct heat, we will include their area and assume it is part of the approximed cylinder's area. Assumed area is then 1.362m^2. Let's take 600 seconds (10 minutes) for the time spent on the ground between tanks filling and aircraft lift-off. We assume the temperature to be constant on the ground and in the tank while heat is transferred. This is correct since the heat energy feeds a phase change in the propellant (vaporization) and not a temperature increase of some sort. ''dQ/dt'' = ''h'' * ''A'' * (''T_ground'' - ''T_prop'') = 0.024382 * 1.362 * (298.15 - 92) = 6.846 J. We can now [[Rocket_Main_Tank#Calculating_evaporation_rate|calculate]] the evaporation rate of the propellant with and without insulation. The heat of vaporization for [[LOX]] is 213 kJ/kg. Without insulation, with this erroneous ''h'' value, the rate of vaporization is 32mg/s, making 19 grammes of LOX evaporated in 600s. That doesn't seem quite right... With a 10mm [[Rocket_Main_Tank#Thermal_insulation_materials|Aerogel]] insulation, the thermal conductivity being 15 mW/m.K, its U-value being k/L = 0.015/0.010 = 1.5 W/m^2.K. ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2], ''Heat transfer in turbulent flow over a flat plate'', is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr varies between 0.73 and 0.75, and the Reynolds number for the characteristic length L of the tank (Re_L) drops from 4.36714e+07 to 6.68455e+06 with altitude. This value of Re confirms that the boundary layer is turbulent. [2] indicates that the Colburn analogy can still be used for Pr up to 10^8, but with some loss of accuracy. We use the following approximation to calculate the average Nusselt number for forced convection in turbulent flows: Nu_L = 0.0366 (Re_L^0.8 - 14500) Pr^1/3 given in [2]. The calculated average Nusselt number for ascent operation is 42353 at sea level and drops to 1037 at 30km altitude. ''h_L'' can then be calculated too, ''h_L'' = Nu_L * ''k''_air / L. It varies from 460.8 on the ground to 9.6 W/m^2.K at 30km altitude. The simple flight path approximation is used to calculate the average ''h'' value for every second of the flight. We also need to calculate the dQ/dt value for each second, since the external temperature changes as the aircraft ascends. The integrated value of heat Q is finally: Q = 1144.12 kJ. '''For [[LOX]], it translates into more than 53kg of oxidizer evaporated during the ascent, when no insulation is used. With an Aerogel insulation layer of 10mm, the mass of evaporated LOX is evaluated to 0.87kg.''' That corresponds to Q = 185.41 kJ. The overall heat transfer coefficient U was calculated using [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#Alternative_Method_.28A_simple_method_for_determining_the_overall_heat_transfer_coefficient.29 this] method. ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 3. Lecture on '''Advanced Heat Transfer''', Ibrahim Sezai, 2011. Distributed as open courseware, [http://opencourses.emu.edu.tr/course/view.php?id=23 here] b301d80860bbf952f3746df14c5fb891e01b49a4 422 421 2012-11-23T04:22:17Z Vincent 1 /* Cases of application */ about the source, luke wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = Nu The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ===The Rayleigh number ''Ra''=== The [https://en.wikipedia.org/wiki/Rayleigh_Number Rayleigh number] is an indicator for natural convection heat transfer. When it is below the critical value for a fluid, heat transfer is primarily in the form of conduction; when it exceeds the critical value, heat transfer is primarily in the form of convection. {{SERVER}}/images/formulas_mirror/rayleigh_number_neg.png α is the [https://en.wikipedia.org/wiki/Thermal_diffusivity thermal diffusivity] {{SERVER}}/images/formulas_mirror/thermal_diffusivity_neg.png ; β is the [https://en.wikipedia.org/wiki/Coefficient_of_thermal_expansion thermal expansion coefficient], for an isobaric process it can be approximated to β = 1/T. ''x'' is the distance from the leading edge. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of cryogenic propellant on the ground and during the ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. The program using the [[Flight_at_high_altitude#Gas_properties_and_altitude|atmospheric model]] and the flight path model to compute the heat transfer during flight will be published freely soon. ===Natural convection for horizontal cryogenic tank=== '''''The results presented in this section are not yet correct. An error has been identified in the calculation of h, but it may not be the only one.''''' On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. {{SERVER}}/images/formulas_mirror/h_natural_conv_external_horiz_cyl_neg.png D is the diameter of the approximated cylinder, in our case we take 0.266m, given by a [[Rocket:First_approximations#Minimum_mass_evaluation|first evaluation]] of the rocket mass program. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C or 298.15 K ground temperature: * the density (rho) is 1.18391 kg/m^3 and at the mean film temperature of -78.07°C or 195.075 K, given the propellant and tank temperature of 92 K: * the viscosity µ is 1.30015e-05 Pa.s (calculated using [http://www.cfd-online.com/Wiki/Sutherland's_law Sutherland's law]) * the kinematic viscosity nu is µ / rho = 1.0982e-5 m^2/s * the specific heat Cp is 1007.68 J/kg.K (approximated using [http://ninova.itu.edu.tr/tr/dersler/ucak-uzay-fakultesi/965/uck-421/ekkaynaklar?g96162 this interpolation]) * the Prandtl number is 0.874462 * the thermal conductivity of air ''k'' is 0.0177375 W/m.K (approximated using [http://physics.tutorvista.com/heat/heat-transfer.html#thermal-conductivity-of-air this interpolation]) * α is thus k/(rho.Cp) = 1.4868e-5 m^2/s * β is approximated to 1/T = 1/195.075 = 5.126e-3 1/K * Ra_D is ((9.80665 * 5.126e-3) / (1.0982e-5 * 1.4868e-5)) * (298.15 - 92) * 0.266^3 = 6.3467e10 * 0.266^3 = 1.1945e9. * finally, ''h'' is 0.0177375/0.266 * (0.6 + (0.387 * 1.1945e9^1/6) / (1 + (0.559/0.874462)^9/16)^8/27)^2 '''= 0.024382 W/m^2.K''' . '''''WARNING: h should bee much higher.''''' We finally have a heat transfer coefficient for air. We can compute the total heat transferred from it, using Newton's law equation [[Heat_transfer#Heat_transfer_coefficient_h|(top)]]. We need to specify the area on which the heat transfer will apply, the approximated cylinder, and the number of seconds during which the heat is transferred. Tanks are a cylinder with two hemispherical end-caps. Since the end-caps still conduct heat, we will include their area and assume it is part of the approximed cylinder's area. Assumed area is then 1.362m^2. Let's take 600 seconds (10 minutes) for the time spent on the ground between tanks filling and aircraft lift-off. We assume the temperature to be constant on the ground and in the tank while heat is transferred. This is correct since the heat energy feeds a phase change in the propellant (vaporization) and not a temperature increase of some sort. ''dQ/dt'' = ''h'' * ''A'' * (''T_ground'' - ''T_prop'') = 0.024382 * 1.362 * (298.15 - 92) = 6.846 J. We can now [[Rocket_Main_Tank#Calculating_evaporation_rate|calculate]] the evaporation rate of the propellant with and without insulation. The heat of vaporization for [[LOX]] is 213 kJ/kg. Without insulation, with this erroneous ''h'' value, the rate of vaporization is 32mg/s, making 19 grammes of LOX evaporated in 600s. That doesn't seem quite right... With a 10mm [[Rocket_Main_Tank#Thermal_insulation_materials|Aerogel]] insulation, the thermal conductivity being 15 mW/m.K, its U-value being k/L = 0.015/0.010 = 1.5 W/m^2.K. ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2], ''Heat transfer in turbulent flow over a flat plate'', is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr varies between 0.73 and 0.75, and the Reynolds number for the characteristic length L of the tank (Re_L) drops from 4.36714e+07 to 6.68455e+06 with altitude. This value of Re confirms that the boundary layer is turbulent. [2] indicates that the Colburn analogy can still be used for Pr up to 10^8, but with some loss of accuracy. We use the following approximation to calculate the average Nusselt number for forced convection in turbulent flows: Nu_L = 0.0366 (Re_L^0.8 - 14500) Pr^1/3 given in [2]. The calculated average Nusselt number for ascent operation is 42353 at sea level and drops to 1037 at 30km altitude. ''h_L'' can then be calculated too, ''h_L'' = Nu_L * ''k''_air / L. It varies from 460.8 on the ground to 9.6 W/m^2.K at 30km altitude. The simple flight path approximation is used to calculate the average ''h'' value for every second of the flight. We also need to calculate the dQ/dt value for each second, since the external temperature changes as the aircraft ascends. The integrated value of heat Q is finally: Q = 1144.12 kJ. '''For [[LOX]], it translates into more than 53kg of oxidizer evaporated during the ascent, when no insulation is used. With an Aerogel insulation layer of 10mm, the mass of evaporated LOX is evaluated to 0.87kg.''' That corresponds to Q = 185.41 kJ. The overall heat transfer coefficient U was calculated using [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#Alternative_Method_.28A_simple_method_for_determining_the_overall_heat_transfer_coefficient.29 this] method. ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 3. Lecture on '''Advanced Heat Transfer''', Ibrahim Sezai, 2011. Distributed as open courseware, [http://opencourses.emu.edu.tr/course/view.php?id=23 here] 193e6dde7bb1021639c143cef4d2871aaeadb00b 423 422 2012-11-23T16:04:17Z Vincent 1 /* Natural convection for horizontal cryogenic tank */ fixing the wrong calculation of h and first real approximation of ground vap rate wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = Nu The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ===The Rayleigh number ''Ra''=== The [https://en.wikipedia.org/wiki/Rayleigh_Number Rayleigh number] is an indicator for natural convection heat transfer. When it is below the critical value for a fluid, heat transfer is primarily in the form of conduction; when it exceeds the critical value, heat transfer is primarily in the form of convection. {{SERVER}}/images/formulas_mirror/rayleigh_number_neg.png α is the [https://en.wikipedia.org/wiki/Thermal_diffusivity thermal diffusivity] {{SERVER}}/images/formulas_mirror/thermal_diffusivity_neg.png ; β is the [https://en.wikipedia.org/wiki/Coefficient_of_thermal_expansion thermal expansion coefficient], for an isobaric process it can be approximated to β = 1/T. ''x'' is the distance from the leading edge. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of cryogenic propellant on the ground and during the ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. The program using the [[Flight_at_high_altitude#Gas_properties_and_altitude|atmospheric model]] and the flight path model to compute the heat transfer during flight will be published freely soon. ===Natural convection for horizontal cryogenic tank=== On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. {{SERVER}}/images/formulas_mirror/h_natural_conv_external_horiz_cyl_neg.png D is the diameter of the approximated cylinder, in our case we take 0.266m, given by a [[Rocket:First_approximations#Minimum_mass_evaluation|first evaluation]] of the rocket mass program. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C or 298.15 K ground temperature: * the density (rho) is 1.18391 kg/m^3 and at the mean film temperature of -78.07°C or 195.075 K, given the propellant and tank temperature of 92 K: * the viscosity µ is 1.30015e-05 Pa.s (calculated using [http://www.cfd-online.com/Wiki/Sutherland's_law Sutherland's law]) * the kinematic viscosity nu is µ / rho = 1.0982e-5 m^2/s * the specific heat Cp is 1007.68 J/kg.K (approximated using [http://ninova.itu.edu.tr/tr/dersler/ucak-uzay-fakultesi/965/uck-421/ekkaynaklar?g96162 this interpolation]) * the Prandtl number is 0.73862 * the thermal conductivity of air ''k'' is 0.0177375 W/m.K (approximated using [http://physics.tutorvista.com/heat/heat-transfer.html#thermal-conductivity-of-air this interpolation]) * α is thus k/(rho.Cp) = 1.4868e-5 m^2/s * β is approximated to 1/T = 1/195.075 = 5.126e-3 1/K * Ra_D is ((9.80665 * 5.126e-3) / (1.0982e-5 * 1.4868e-5)) * (298.15 - 92) * 0.266^3 = 6.3467e10 * 0.266^3 = 1.24931e+9. * finally, ''h'' is 0.0177375/0.266 * (0.6 + (0.387 * 1.24931e+9^1/6) / (1 + (0.559/0.73862)^9/16)^8/27)^2 '''= 8.20533 W/m^2.K''' We finally have a heat transfer coefficient for air. We can compute the total heat transferred from it, using Newton's law equation [[Heat_transfer#Heat_transfer_coefficient_h|(top)]]. We need to specify the area on which the heat transfer will apply, the approximated cylinder, and the number of seconds during which the heat is transferred. Tanks are a cylinder with two hemispherical end-caps. Since the end-caps still conduct heat, we will include their area and assume it is part of the approximed cylinder's area. Assumed area is then 1.3826m^2. Let's take 600 seconds (10 minutes) for the time spent on the ground between tanks filling and aircraft lift-off. We assume the temperature to be constant on the ground and in the tank while heat is transferred. This is correct since the heat energy feeds a phase change in the propellant (vaporization) and not a temperature increase of some sort. ''dQ/dt'' = ''h'' * ''A'' * (''T_ground'' - ''T_prop'') = 8.20533 * 1.3826 * (298.15 - 92) = 2338.73 J/s. We can now [[Rocket_Main_Tank#Calculating_evaporation_rate|calculate]] the evaporation rate of the propellant with and without insulation. The heat of vaporization for [[LOX]] is 213 kJ/kg. Without insulation, the vaporization rate of LOX is 11g/s, making 6.59kg of LOX evaporated in 600s. With a 10mm [[Rocket_Main_Tank#Thermal_insulation_materials|Aerogel]] insulation, the thermal conductivity being 15 mW/m.K, its U-value being k/L = 0.015/0.010 = 1.5 W/m^2.K, the vaporization rate is 919.5mg/s. For 10 minutes, it goes down to 0.552kg (instead of 6.59 without insulation). ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2], ''Heat transfer in turbulent flow over a flat plate'', is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr varies between 0.73 and 0.75, and the Reynolds number for the characteristic length L of the tank (Re_L) drops from 4.36714e+07 to 6.68455e+06 with altitude. This value of Re confirms that the boundary layer is turbulent. [2] indicates that the Colburn analogy can still be used for Pr up to 10^8, but with some loss of accuracy. We use the following approximation to calculate the average Nusselt number for forced convection in turbulent flows: Nu_L = 0.0366 (Re_L^0.8 - 14500) Pr^1/3 given in [2]. The calculated average Nusselt number for ascent operation is 42353 at sea level and drops to 1037 at 30km altitude. ''h_L'' can then be calculated too, ''h_L'' = Nu_L * ''k''_air / L. It varies from 460.8 on the ground to 9.6 W/m^2.K at 30km altitude. The simple flight path approximation is used to calculate the average ''h'' value for every second of the flight. We also need to calculate the dQ/dt value for each second, since the external temperature changes as the aircraft ascends. The integrated value of heat Q is finally: Q = 1144.12 kJ. '''For [[LOX]], it translates into more than 53kg of oxidizer evaporated during the ascent, when no insulation is used. With an Aerogel insulation layer of 10mm, the mass of evaporated LOX is evaluated to 0.87kg.''' That corresponds to Q = 185.41 kJ. The overall heat transfer coefficient U was calculated using [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#Alternative_Method_.28A_simple_method_for_determining_the_overall_heat_transfer_coefficient.29 this] method. ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 3. Lecture on '''Advanced Heat Transfer''', Ibrahim Sezai, 2011. Distributed as open courseware, [http://opencourses.emu.edu.tr/course/view.php?id=23 here] 55e0ef74e56d0e650c941d55dc20871e5e7729c6 424 423 2012-11-23T16:20:39Z Vincent 1 /* Cases of application */ fixing the flight evaluation of vap, it was calculated with the 5mm insulation instead of 10mm and wrong tank size, warning msgs wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = Nu The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ===The Rayleigh number ''Ra''=== The [https://en.wikipedia.org/wiki/Rayleigh_Number Rayleigh number] is an indicator for natural convection heat transfer. When it is below the critical value for a fluid, heat transfer is primarily in the form of conduction; when it exceeds the critical value, heat transfer is primarily in the form of convection. {{SERVER}}/images/formulas_mirror/rayleigh_number_neg.png α is the [https://en.wikipedia.org/wiki/Thermal_diffusivity thermal diffusivity] {{SERVER}}/images/formulas_mirror/thermal_diffusivity_neg.png ; β is the [https://en.wikipedia.org/wiki/Coefficient_of_thermal_expansion thermal expansion coefficient], for an isobaric process it can be approximated to β = 1/T. ''x'' is the distance from the leading edge. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of cryogenic propellant on the ground and during the ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. The program using the [[Flight_at_high_altitude#Gas_properties_and_altitude|atmospheric model]] and the flight path model to compute the heat transfer during flight will be published freely soon. '''''Warning:''''' ''the values presented below are highly dependent on the dimensions of the considered tank (diameter: 0.27m, length: 1.63m, surface: 1.38261m^2), which was given by the [[Rocket:First_approximations#Minimum_mass_evaluation|first approximation]] of the rocket mass program, the climb rates and flight profile (currently based on [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2]'s rate), and the temperature of the propellant (here LOX at 92K).'' '''''Warning:''''' ''the method and results presented here are based on approximations and assumptions, and it may even have not been done in the proper way. Please validate the method used if you have some knowledge on convective heat transfer.'' ===Natural convection for horizontal cryogenic tank=== On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. {{SERVER}}/images/formulas_mirror/h_natural_conv_external_horiz_cyl_neg.png D is the diameter of the approximated cylinder, in our case we take 0.27m, given by a [[Rocket:First_approximations#Minimum_mass_evaluation|first evaluation]] of the rocket mass program. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C or 298.15 K ground temperature: * the density (rho) is 1.18391 kg/m^3 and at the mean film temperature of -78.07°C or 195.075 K, given the propellant and tank temperature of 92 K: * the viscosity µ is 1.30015e-05 Pa.s (calculated using [http://www.cfd-online.com/Wiki/Sutherland's_law Sutherland's law]) * the kinematic viscosity nu is µ / rho = 1.0982e-5 m^2/s * the specific heat Cp is 1007.68 J/kg.K (approximated using [http://ninova.itu.edu.tr/tr/dersler/ucak-uzay-fakultesi/965/uck-421/ekkaynaklar?g96162 this interpolation]) * the Prandtl number is 0.73862 * the thermal conductivity of air ''k'' is 0.0177375 W/m.K (approximated using [http://physics.tutorvista.com/heat/heat-transfer.html#thermal-conductivity-of-air this interpolation]) * α is thus k/(rho.Cp) = 1.4868e-5 m^2/s * β is approximated to 1/T = 1/195.075 = 5.126e-3 1/K * Ra_D is ((9.80665 * 5.126e-3) / (1.0982e-5 * 1.4868e-5)) * (298.15 - 92) * 0.27^3 = 6.3467e10 * 0.27^3 = 1.24931e+9. * finally, ''h'' is 0.0177375/0.27 * (0.6 + (0.387 * 1.24931e+9^1/6) / (1 + (0.559/0.73862)^9/16)^8/27)^2 '''= 8.20533 W/m^2.K''' We finally have a heat transfer coefficient for air. We can compute the total heat transferred from it, using Newton's law equation [[Heat_transfer#Heat_transfer_coefficient_h|(top)]]. We need to specify the area on which the heat transfer will apply, the approximated cylinder, and the number of seconds during which the heat is transferred. Tanks are a cylinder with two hemispherical end-caps. Since the end-caps still conduct heat, we will include their area and assume it is part of the approximed cylinder's area. Assumed area is then 1.3826m^2. Let's take 600 seconds (10 minutes) for the time spent on the ground between tanks filling and aircraft lift-off. We assume the temperature to be constant on the ground and in the tank while heat is transferred. This is correct since the heat energy feeds a phase change in the propellant (vaporization) and not a temperature increase of some sort. ''dQ/dt'' = ''h'' * ''A'' * (''T_ground'' - ''T_prop'') = 8.20533 * 1.3826 * (298.15 - 92) = 2338.73 J/s. We can now [[Rocket_Main_Tank#Calculating_evaporation_rate|calculate]] the evaporation rate of the propellant with and without insulation. The heat of vaporization for [[LOX]] is 213 kJ/kg. Without insulation, the vaporization rate of LOX is 11g/s, making 6.59kg of LOX evaporated in 600s. With a 10mm [[Rocket_Main_Tank#Thermal_insulation_materials|Aerogel]] insulation, the thermal conductivity being 15 mW/m.K, its U-value being k/L = 0.015/0.010 = 1.5 W/m^2.K, the vaporization rate is 919.5mg/s. For 10 minutes, it goes down to 0.552kg (instead of 6.59 without insulation). The overall heat transfer coefficient U was calculated using [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#Alternative_Method_.28A_simple_method_for_determining_the_overall_heat_transfer_coefficient.29 this] method. ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2], ''Heat transfer in turbulent flow over a flat plate'', is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr varies between 0.73 and 0.75, and the Reynolds number for the characteristic length L of the tank (Re_L) drops from 4.36714e+07 to 6.68455e+06 with altitude. This value of Re confirms that the boundary layer is turbulent. [2] indicates that the Colburn analogy can still be used for Pr up to 10^8, but with some loss of accuracy. We use the following approximation to calculate the average Nusselt number for forced convection in turbulent flows: Nu_L = 0.0366 (Re_L^0.8 - 14500) Pr^1/3 given in [2]. The calculated average Nusselt number for ascent operation is 42353 at sea level and drops to 1037 at 30km altitude. ''h_L'' can then be calculated too, ''h_L'' = Nu_L * ''k''_air / L. It varies from 460.8 on the ground to 9.6 W/m^2.K at 30km altitude. The simple flight path approximation is used to calculate the average ''h'' value for every second of the flight. We also need to calculate the dQ/dt value for each second, since the external temperature changes as the aircraft ascends. The integrated value of heat Q is finally: Q = 1710.94 kJ. '''For [[LOX]], it translates into more than 80kg of oxidizer evaporated during the ascent, when no insulation is used. With an Aerogel insulation layer of 10mm, the mass of evaporated LOX is evaluated to 0.666kg.''' That corresponds to Q = 141.89 kJ. The overall heat transfer coefficient U was calculated using [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#Alternative_Method_.28A_simple_method_for_determining_the_overall_heat_transfer_coefficient.29 this] method. ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 3. Lecture on '''Advanced Heat Transfer''', Ibrahim Sezai, 2011. Distributed as open courseware, [http://opencourses.emu.edu.tr/course/view.php?id=23 here] beba689199c34959b5e1bc18482764c8f97c2a62 428 424 2012-11-24T02:10:19Z Vincent 1 /* Cases of application */ source code link wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = Nu The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ===The Rayleigh number ''Ra''=== The [https://en.wikipedia.org/wiki/Rayleigh_Number Rayleigh number] is an indicator for natural convection heat transfer. When it is below the critical value for a fluid, heat transfer is primarily in the form of conduction; when it exceeds the critical value, heat transfer is primarily in the form of convection. {{SERVER}}/images/formulas_mirror/rayleigh_number_neg.png α is the [https://en.wikipedia.org/wiki/Thermal_diffusivity thermal diffusivity] {{SERVER}}/images/formulas_mirror/thermal_diffusivity_neg.png ; β is the [https://en.wikipedia.org/wiki/Coefficient_of_thermal_expansion thermal expansion coefficient], for an isobaric process it can be approximated to β = 1/T. ''x'' is the distance from the leading edge. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of cryogenic propellant on the ground and during the ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. The program using the [[Flight_at_high_altitude#Gas_properties_and_altitude|atmospheric model]] and the flight path model to compute the heat transfer during flight is available freely here: [[File:Heat_transfer_to_rocket_tank.c]]. '''''Warning:''''' ''the values presented below are highly dependent on the dimensions of the considered tank (diameter: 0.27m, length: 1.63m, surface: 1.38261m^2), which was given by the [[Rocket:First_approximations#Minimum_mass_evaluation|first approximation]] of the rocket mass program, the climb rates and flight profile (currently based on [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2]'s rate), and the temperature of the propellant (here LOX at 92K).'' '''''Warning:''''' ''the method and results presented here are based on approximations and assumptions, and it may even have not been done in the proper way. Please validate the method used if you have some knowledge on convective heat transfer.'' ===Natural convection for horizontal cryogenic tank=== On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. {{SERVER}}/images/formulas_mirror/h_natural_conv_external_horiz_cyl_neg.png D is the diameter of the approximated cylinder, in our case we take 0.27m, given by a [[Rocket:First_approximations#Minimum_mass_evaluation|first evaluation]] of the rocket mass program. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C or 298.15 K ground temperature: * the density (rho) is 1.18391 kg/m^3 and at the mean film temperature of -78.07°C or 195.075 K, given the propellant and tank temperature of 92 K: * the viscosity µ is 1.30015e-05 Pa.s (calculated using [http://www.cfd-online.com/Wiki/Sutherland's_law Sutherland's law]) * the kinematic viscosity nu is µ / rho = 1.0982e-5 m^2/s * the specific heat Cp is 1007.68 J/kg.K (approximated using [http://ninova.itu.edu.tr/tr/dersler/ucak-uzay-fakultesi/965/uck-421/ekkaynaklar?g96162 this interpolation]) * the Prandtl number is 0.73862 * the thermal conductivity of air ''k'' is 0.0177375 W/m.K (approximated using [http://physics.tutorvista.com/heat/heat-transfer.html#thermal-conductivity-of-air this interpolation]) * α is thus k/(rho.Cp) = 1.4868e-5 m^2/s * β is approximated to 1/T = 1/195.075 = 5.126e-3 1/K * Ra_D is ((9.80665 * 5.126e-3) / (1.0982e-5 * 1.4868e-5)) * (298.15 - 92) * 0.27^3 = 6.3467e10 * 0.27^3 = 1.24931e+9. * finally, ''h'' is 0.0177375/0.27 * (0.6 + (0.387 * 1.24931e+9^1/6) / (1 + (0.559/0.73862)^9/16)^8/27)^2 '''= 8.20533 W/m^2.K''' We finally have a heat transfer coefficient for air. We can compute the total heat transferred from it, using Newton's law equation [[Heat_transfer#Heat_transfer_coefficient_h|(top)]]. We need to specify the area on which the heat transfer will apply, the approximated cylinder, and the number of seconds during which the heat is transferred. Tanks are a cylinder with two hemispherical end-caps. Since the end-caps still conduct heat, we will include their area and assume it is part of the approximed cylinder's area. Assumed area is then 1.3826m^2. Let's take 600 seconds (10 minutes) for the time spent on the ground between tanks filling and aircraft lift-off. We assume the temperature to be constant on the ground and in the tank while heat is transferred. This is correct since the heat energy feeds a phase change in the propellant (vaporization) and not a temperature increase of some sort. ''dQ/dt'' = ''h'' * ''A'' * (''T_ground'' - ''T_prop'') = 8.20533 * 1.3826 * (298.15 - 92) = 2338.73 J/s. We can now [[Rocket_Main_Tank#Calculating_evaporation_rate|calculate]] the evaporation rate of the propellant with and without insulation. The heat of vaporization for [[LOX]] is 213 kJ/kg. Without insulation, the vaporization rate of LOX is 11g/s, making 6.59kg of LOX evaporated in 600s. With a 10mm [[Rocket_Main_Tank#Thermal_insulation_materials|Aerogel]] insulation, the thermal conductivity being 15 mW/m.K, its U-value being k/L = 0.015/0.010 = 1.5 W/m^2.K, the vaporization rate is 919.5mg/s. For 10 minutes, it goes down to 0.552kg (instead of 6.59 without insulation). The overall heat transfer coefficient U was calculated using [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#Alternative_Method_.28A_simple_method_for_determining_the_overall_heat_transfer_coefficient.29 this] method. ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2], ''Heat transfer in turbulent flow over a flat plate'', is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr varies between 0.73 and 0.75, and the Reynolds number for the characteristic length L of the tank (Re_L) drops from 4.36714e+07 to 6.68455e+06 with altitude. This value of Re confirms that the boundary layer is turbulent. [2] indicates that the Colburn analogy can still be used for Pr up to 10^8, but with some loss of accuracy. We use the following approximation to calculate the average Nusselt number for forced convection in turbulent flows: Nu_L = 0.0366 (Re_L^0.8 - 14500) Pr^1/3 given in [2]. The calculated average Nusselt number for ascent operation is 42353 at sea level and drops to 1037 at 30km altitude. ''h_L'' can then be calculated too, ''h_L'' = Nu_L * ''k''_air / L. It varies from 460.8 on the ground to 9.6 W/m^2.K at 30km altitude. The simple flight path approximation is used to calculate the average ''h'' value for every second of the flight. We also need to calculate the dQ/dt value for each second, since the external temperature changes as the aircraft ascends. The integrated value of heat Q is finally: Q = 1710.94 kJ. '''For [[LOX]], it translates into more than 80kg of oxidizer evaporated during the ascent, when no insulation is used. With an Aerogel insulation layer of 10mm, the mass of evaporated LOX is evaluated to 0.666kg.''' That corresponds to Q = 141.89 kJ. The overall heat transfer coefficient U was calculated using [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#Alternative_Method_.28A_simple_method_for_determining_the_overall_heat_transfer_coefficient.29 this] method. ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 3. Lecture on '''Advanced Heat Transfer''', Ibrahim Sezai, 2011. Distributed as open courseware, [http://opencourses.emu.edu.tr/course/view.php?id=23 here] 45359a1376492972dec4f9e61f5a59722bebbfa0 0 56 408 347 2012-11-20T22:42:54Z Vincent 1 /* Resources on physics related to aerodynamics */ heat transfer link wikitext text/x-wiki =Resources on physics related to aerodynamics= The [https://en.wikipedia.org/wiki/List_of_elementary_physics_formulae List of elementary physics formulae] on wikipedia is useful. A page is dedicated to [[heat transfer]]. ==List of variables== {| border="1" class="wikitable" !Variable !Meaning !Unit (SI) |- | γ (gamma) | [https://en.wikipedia.org/wiki/Surface_tension Surface tension] or [https://en.wikipedia.org/wiki/Heat_capacity_ratio Heat capacity ratio] (adiabatic process in thermodynamics) | N.m^-1 (Newton per meter) |- | μ (mu) or η (eta) | [https://en.wikipedia.org/wiki/Viscosity Viscosity] | Pa·s (Pascal second) or P (Poise, 1 Poise is 0.1 Pa.s) |- | ρ (rho) | [https://en.wikipedia.org/wiki/Density Density] | kg.m^-3 (kg per cubic meter) |- | C, C_p, C_V | [https://en.wikipedia.org/wiki/Heat_capacity#Metrology Heat capacity], general, at constant pressure, at constant volume. | J.K^-1 (Joule per Kelvin) |- | G | [https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] | J (Joule) |- | H | [https://en.wikipedia.org/wiki/Enthalpy Enthalpy]: total energy of a thermodynamic system. | J (Joule) |- | ΔH_vap or L | [https://en.wikipedia.org/wiki/Vaporization_heat Vaporization heat] or [https://en.wikipedia.org/wiki/Latent_heat Latent heat of vaporization]: energy required to vaporize a mole of liquid at a given temperature. | J.mol^-1 (Joule per mole) |- | M | [https://en.wikipedia.org/wiki/Mach_number Mach number] | no unit |- | Q | Amount of [https://en.wikipedia.org/wiki/Heat Heat] | J (Joule) |- | T | Temperature. T₀ or T_t is the [https://en.wikipedia.org/wiki/Stagnation_temperature stagnation temperature]. | K (Kelvin) |- | S | [https://en.wikipedia.org/wiki/Entropy Entropy] | J.K^-1 (Joule per Kelvin) |- | U | [https://en.wikipedia.org/wiki/Internal_energy Internal energy] of a system (see first law of Thermodynamics below) | J (Joule) |- | V | Volume | m³ (cubic meter) |- | W | [https://en.wikipedia.org/wiki/Work_(thermodynamics) Work]: mechanical constraints on the system. | J (Joule) |- | a | [https://en.wikipedia.org/wiki/Speed_of_sound Speed of sound] in medium (used to calculate Mach number) | m.s^-1 |- | c | Velocity of a flow in thermodynamics, also noted V; generally noted u in fluid dynamics. | m.s^-1 |- | n | Quantity of matter | mol (mole) |- | p | Pressure. p_t is the [https://en.wikipedia.org/wiki/Stagnation_pressure stagnation pressure]. | Pa (Pascal) |- | p | [https://en.wikipedia.org/wiki/Momentum Momentum] p = m*v, with m the mass and v the velocity, not to be confused with volume. | kg.m.s^-1 |} ==List of constants== {| border="1" class="wikitable" !Constant !Meaning !Value !Unit (SI) |- | N_A or N | [https://en.wikipedia.org/wiki/Avogadro_constant Avogadro constant], number of atoms or molecules in a mole. | 6.02214129.10²³ | mol^-1 |- | R | [https://en.wikipedia.org/wiki/Gas_constant ideal gas constant] | 8.3144621 | J.K⁻¹.mol⁻¹ |- | G | [https://en.wikipedia.org/wiki/Gravitational_constant Gravitational constant] | 6.674 | m³.kg^-1.s^-2 |- | k_B or k | [https://en.wikipedia.org/wiki/Boltzmann_constant Boltzmann constant], gas constant R divided by Avogadro number. | 1.3806488.10^-23 | J.K^-1 |} ==List of equations== {| border="1" class="wikitable" !Equation !Name !Meaning |- |{{SERVER}}/images/formulas_mirror/pvnrtk_neg.png |Ideal gas equation |Relation between properties of an ideal gas ([https://en.wikipedia.org/wiki/State_equation state equation]). k is k_B. |- |{{SERVER}}/images/formulas_mirror/clausius-clapeyron_neg.png |[https://en.wikipedia.org/wiki/Clausius%E2%80%93Clapeyron_relation#Ideal_gas_approximation_at_low_temperatures Clausius-Clapeyron relation] |Relation between the pressure, latent heat of vaporization and temperature of a vapour at two temperatures (approximation, at low temperatures). |- |{{SERVER}}/images/formulas_mirror/QeqmL_neg.png |Heat at [https://en.wikipedia.org/wiki/Latent_heat#Specific_latent_heat state change] for an ideal gas. |The heat required to change the state of a some matter, L being the latent heat. Delta H equals Q only when pressure is constant (isobaric). |- |{{SERVER}}/images/formulas_mirror/dUeqdQmindW_neg.png |[https://en.wikipedia.org/wiki/First_law_of_thermodynamics First law of thermodynamics] |Variations of internal energy of a system between two states is the sum of the received heat and work (minus the ''given'' work). |- |{{SERVER}}/images/formulas_mirror/enthalpy_neg.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] |Total amount of energy of a system, defined as the sum of the internal energy U of the system and pressure * volume at the boundary of the system and its environment. |- |{{SERVER}}/images/formulas_mirror/workExpand_neg.png |Work of gas expansion. |Work done by expanding an ideal gas. |- |{{SERVER}}/images/formulas_mirror/entropy_dueqtdsmpdv_neg.png |[https://en.wikipedia.org/wiki/Internal_energy Internal energy] change related to [https://en.wikipedia.org/wiki/Entropy entropy] |Internal energy related to entropy variation for a closed system in thermal equilibrium ([https://en.wikipedia.org/wiki/Fundamental_thermodynamic_relation fundamental thermodynamic relation]). |- |{{SERVER}}/images/formulas_mirror/dheqtds_neg.png |[https://en.wikipedia.org/wiki/Enthalpy Enthalpy] change |Enthalpy change depending on entropy and pressure changes, equation created from the mix of the basic ones above. |- |ΔS_universe = ΔS_surroundings + ΔS_system |Entropy variation as a whole. |Entropy variation of a system is generally compensated by the inverse variation of the surroundings, not including losses. |- |{{SERVER}}/images/formulas_mirror/dS_neg.png |[https://en.wikipedia.org/wiki/Second_law_of_thermodynamics Second law of thermodynamics] |A change in the entropy of a system is the infinitesimal transfer of heat to a closed system driving a reversible process, divided by the equilibrium temperature of the system. |- |{{SERVER}}/images/formulas_mirror/gibbs_neg.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] / Free enthalpy |Useful work obtainable from a system at '''isobaric''' and '''isothermal''' conditions. Since H is U + pV, it can be replaced in the equation, making G = H - TS. |- |{{SERVER}}/images/formulas_mirror/deltaG_neg.png |[https://en.wikipedia.org/wiki/Gibbs_free_energy Gibbs free energy] variation. |If ΔG < 0, the system's transformation can be spontaneous, if ΔG = 0 the transformation is inversible and the system is in an equilibrium state, if ΔG > 0 it can't be spontaneous. |- |{{SERVER}}/images/formulas_mirror/density_ideal_neg.png |[https://en.wikipedia.org/wiki/Density#Changes_of_density Density] of an ideal gas. |M is molar mass. This means that the density of an ideal gas can be doubled by doubling the pressure, or by halving the absolute temperature. |} 376511d987dfddecca79f313dedeff12d4078b17 0 16 409 330 2012-11-20T22:44:04Z Vincent 1 /* Resources */ heat transfer link wikitext text/x-wiki =Resources= This page gathers available documentation on the numerous subjects linked the project, including rocket science, turbofans, aerodynamics, amateur rocketry and SSTO. An [[Aero formulas|internal page]] has been created containing useful formulas related to thermodynamics and aerodynamics. Another page contains all [[heat transfer]]-related material. ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/ amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * An important nasaspaceflight [http://forum.nasaspaceflight.com/index.php?topic=2847.0 forum thread] where ideas on micro-rocketry to orbit are discussed and shared. * A kind of spin-off of the above thread is the [http://orbitalaspirations.blogspot.com/ orbital aspirations weblog]. It was recently created by Ed LeBouthillier and is already filled with lots of information on scaling down rocket equations to micro-rocketry, SSTO, reference papers, news of the domain and so on. * Another blog talking about SSTO and air-to-orbit: [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html selenian boondocks]. * Another blog is dedicated to SSTO: [http://exoscientist.blogspot.fr/ Polymath/exoscientist]. * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] by Klaus Hünecke, 1997. * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ===Other: Thermodynamics, Aerodynamics, general physics=== * [http://www.desktop.aero/appliedaero/preface/welcome.html Applied Aerodynamics: A Digital Textbook] by Desktop Aeronautics, Inc. 2007. * [http://www.rshanthini.com/ThermoBook.htm Thermodynamics for Beginners, with worked examples] by R. Shanthini. 2006. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. 3a3f01f305fe5bb82dc2e4c6f5cacb221dff1308 439 409 2012-11-29T03:17:48Z Vincent 1 /* Web pages */ adding gravity loss wikitext text/x-wiki =Resources= This page gathers available documentation on the numerous subjects linked the project, including rocket science, turbofans, aerodynamics, amateur rocketry and SSTO. An [[Aero formulas|internal page]] has been created containing useful formulas related to thermodynamics and aerodynamics. Another page contains all [[heat transfer]]-related material. ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/ amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * An important nasaspaceflight [http://forum.nasaspaceflight.com/index.php?topic=2847.0 forum thread] where ideas on micro-rocketry to orbit are discussed and shared. * A kind of spin-off of the above thread is the [http://orbitalaspirations.blogspot.com/ Orbital Aspirations weblog]. It was recently created by Ed LeBouthillier and is already filled with lots of information on scaling down rocket equations to micro-rocketry, SSTO, reference papers, news of the domain and so on. * Another blog talking about SSTO and air-to-orbit: [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Selenian Boondocks]. * Another blog is dedicated to SSTO: [http://exoscientist.blogspot.fr/ Polymath/Exoscientist]. * A scientist blog with some occasional rocketry: [http://www.gravityloss.com/ Gravity Loss]. * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] by Klaus Hünecke, 1997. * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ===Other: Thermodynamics, Aerodynamics, general physics=== * [http://www.desktop.aero/appliedaero/preface/welcome.html Applied Aerodynamics: A Digital Textbook] by Desktop Aeronautics, Inc. 2007. * [http://www.rshanthini.com/ThermoBook.htm Thermodynamics for Beginners, with worked examples] by R. Shanthini. 2006. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. 55ba1a067bc9fa170d3a9d1e49913ef7eaa5b41b 6 76 426 2012-11-24T02:01:08Z Vincent 1 Program calculating the heat transferred to a cryogenic tank (typically LOX) for balloon- or air-to-orbit vehicles. Two heat transfer approximations are done. * when the vehicle is waiting on the ground after tank filling and before it lifts-off from the wikitext text/x-wiki Program calculating the heat transferred to a cryogenic tank (typically LOX) for balloon- or air-to-orbit vehicles. Two heat transfer approximations are done. * when the vehicle is waiting on the ground after tank filling and before it lifts-off from the ground * during the ascent flight, following configurable climb rates and the ISA atmospheric model already presented here: [[File:ISA_atmospheric_model.c]]. The vaporization rate is given for ground operation, and the overall heat and vaporized mass is given for both. The atmospheric model is at the beginning of the file, other parameters for flight and tanks are right after it. Compile with: <code><gcc -o heat_transfer_to_rocket_tank -Wall heat_transfer_to_rocket_tank.c -lm/code> 4beeed91b58f26bab6cbc58a7d63b25ac0121dee 427 426 2012-11-24T02:01:56Z Vincent 1 fixing code block wikitext text/x-wiki Program calculating the heat transferred to a cryogenic tank (typically LOX) for balloon- or air-to-orbit vehicles. Two heat transfer approximations are done. * when the vehicle is waiting on the ground after tank filling and before it lifts-off from the ground * during the ascent flight, following configurable climb rates and the ISA atmospheric model already presented here: [[File:ISA_atmospheric_model.c]]. The vaporization rate is given for ground operation, and the overall heat and vaporized mass is given for both. The atmospheric model is at the beginning of the file, other parameters for flight and tanks are right after it. Compile with: <code>gcc -o heat_transfer_to_rocket_tank -Wall heat_transfer_to_rocket_tank.c -lm</code> 80530de2bf3631554de197a4698674459a59f353 0 61 429 395 2012-11-25T00:32:27Z Vincent 1 /* Delta V achievement */ text update wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Since a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, an even more powerful engine is required and this is an iterative process. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the wall's thickness and material are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. It is, if losses are not taken into account. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * '''DRY MASS: 12.8552 kg''', WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg '''So here we are. 142kg to carry up to 30km altitude.''' The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten, as well as telemetry equipment. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. Rocket parameters are 15.2 mass ratio (9.0 Delta V, 340 constant I_sp), propellants are LOX and E85 at stoichiometric ratio (no venting, no tank insulation), full aluminium structure. Other parameters are the same than above, and the software used was version 1 (see updates here: [[File:Rocket_mass.c]]). [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> 1e4f1dadde1f1abeb672179d0cf12c2ed54c7ef1 430 429 2012-11-25T02:54:18Z Vincent 1 /* Rocket engine */ more about the two thrust-to-weight ratios wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of aerodynamic drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's Pegasus system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s, and it doesn't change much for an altitude ±100km. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine. To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. The engine's thrust-to-weight ratio will probably not be higher than 100. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. This value has a large impact on the thrust and consequently the mass of the engine, and a more precise estimation has to be done as function of the propellant burn rate and the pitch angle of the rocket. For example, with a 1.25 ratio, a 15 degree angle and an initial (release) velocity of 277 m/s, '''assuming constant vehicle mass and thrust of 1.25 times g''', the rocket can travel 24s before losing its vertical velocity, and during this time it climbs 892 m. In reality, 24s after ignition, the thrust has much increased and it may be sufficient to sustain the climb. That's what has to be evaluated, in order to fix the initial pitch angle of the rocket and the initial thrust. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * '''DRY MASS: 12.8552 kg''', WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg '''So here we are. 142kg to carry up to 30km altitude.''' The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten, as well as telemetry equipment. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. Rocket parameters are 15.2 mass ratio (9.0 Delta V, 340 constant I_sp), propellants are LOX and E85 at stoichiometric ratio (no venting, no tank insulation), full aluminium structure. Other parameters are the same than above, and the software used was version 1 (see updates here: [[File:Rocket_mass.c]]). [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> f19d2ec33df4c9b4d893136108139d008c6a12e3 432 430 2012-11-26T23:32:43Z Vincent 1 /* Delta V requirements */ more about gravity drag wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph), but it's initial thrust-to-weight ratio is between 2 and 3. The gravity drag is reduced when the thrust is higher because the longer the rocket is in powered flight, the longer it has to overcome the gravity acceleration, as explained [https://en.wikipedia.org/wiki/Gravity_drag here]. Consequently, a higher thrust-to-weight ratio will result in less gravity drag. The fact that the engine is heavier to provide the additional thrust is still better than having a larger Delta-V to achieve and carrying more propellant. Additionally, the gravity drag can be partly compensated by small supersonic wings on the rocket's fuselage, like on the Pegasus. Since we are in an SSTO perspective, we have to assess if the additional mass of these wings is a bad thing, and also if they can be jettisoned. * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine. To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. The engine's thrust-to-weight ratio will probably not be higher than 100. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. This value has a large impact on the thrust and consequently the mass of the engine, and a more precise estimation has to be done as function of the propellant burn rate and the pitch angle of the rocket. For example, with a 1.25 ratio, a 15 degree angle and an initial (release) velocity of 277 m/s, '''assuming constant vehicle mass and thrust of 1.25 times g''', the rocket can travel 24s before losing its vertical velocity, and during this time it climbs 892 m. In reality, 24s after ignition, the thrust has much increased and it may be sufficient to sustain the climb. That's what has to be evaluated, in order to fix the initial pitch angle of the rocket and the initial thrust. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * '''DRY MASS: 12.8552 kg''', WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg '''So here we are. 142kg to carry up to 30km altitude.''' The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten, as well as telemetry equipment. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. Rocket parameters are 15.2 mass ratio (9.0 Delta V, 340 constant I_sp), propellants are LOX and E85 at stoichiometric ratio (no venting, no tank insulation), full aluminium structure. Other parameters are the same than above, and the software used was version 1 (see updates here: [[File:Rocket_mass.c]]). [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> 2a6e58de990ceb57deb1a375b31548af4c4d1985 434 432 2012-11-27T03:20:46Z Vincent 1 /* Rocket engine */ overall thrust-to-weight: graph and text wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics atmospheric drag] and climb against gravity ([https://en.wikipedia.org/wiki/Gravity_drag gravity drag]) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. ** Atmospheric drag should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ** Gravity drag is about 1100m/s to 1500m/s with ground launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph), but it's initial thrust-to-weight ratio is between 2 and 3. The gravity drag is reduced when the thrust is higher because the longer the rocket is in powered flight, the longer it has to overcome the gravity acceleration, as explained [https://en.wikipedia.org/wiki/Gravity_drag here]. Consequently, a higher thrust-to-weight ratio will result in less gravity drag. The fact that the engine is heavier to provide the additional thrust is still better than having a larger Delta-V to achieve and carrying more propellant. Additionally, the gravity drag can be partly compensated by small supersonic wings on the rocket's fuselage, like on the Pegasus. Since we are in an SSTO perspective, we have to assess if the additional mass of these wings is a bad thing, and also if they can be jettisoned. * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned horizontal release velocity still unknown, possibly 1.5M ±0.5M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is relative to the centre of the Earth and not to the surface. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine. To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. The engine's thrust-to-weight ratio will probably not be higher than 100. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. [[File:Rocket_ignition.png|right]] The rocket release velocity will be the same than the aircraft's, probably meaning a pitch of zero and the rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic, because it would apply very large stress on the vehicle. Moreover, even if it is enough to compensate gravity, it is not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (which not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual, since ground- or balloon-launched rockets start by gaining altitude before orbital speed. These simple examples show how important it is to use the atmosphere to climb. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * '''DRY MASS: 12.8552 kg''', WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg '''So here we are. 142kg to carry up to 30km altitude.''' The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten, as well as telemetry equipment. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. Rocket parameters are 15.2 mass ratio (9.0 Delta V, 340 constant I_sp), propellants are LOX and E85 at stoichiometric ratio (no venting, no tank insulation), full aluminium structure. Other parameters are the same than above, and the software used was version 1 (see updates here: [[File:Rocket_mass.c]]). [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> 2b335d24a42a3b174037bed5cd374f81de0613b3 435 434 2012-11-27T03:47:42Z Vincent 1 fixes along the page wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face atmospheric drag and climb against gravity (gravity drag) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. ** [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ** [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph), but its overall thrust-to-weight ratio is between 2 and 3. The gravity drag is reduced when the thrust is higher because the longer the rocket is in powered flight, the longer it has to overcome the gravity acceleration, as explained [https://en.wikipedia.org/wiki/Gravity_drag here]. Consequently, a higher overall thrust-to-weight ratio will result in less gravity drag. The fact that the engine is heavier to provide the additional thrust is still better than having a larger Delta-V to achieve and carrying more propellant (see the [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] below). Additionally, the gravity drag [[Rocket:First_approximations#Rocket_engine|should]] be partly compensated by small wings on the rocket's fuselage, like on the Pegasus. Since we are in an SSTO perspective, we have to assess how bad the additional mass of these wings is, and also if they can be deployed and jettisoned easily. * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, but do we assume that the air is moving at the same speed than the ground? We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. [[File:Rocket_ignition.png|right]] The rocket release velocity will be the same than the aircraft's, probably meaning a pitch of zero and the rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic, because it would apply very large stress on the vehicle. Moreover, even if it is enough to compensate gravity, it is not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (which not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual, since ground- or balloon-launched rockets start by gaining altitude before orbital speed. These simple examples show how important it is to use the atmosphere to climb. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * '''DRY MASS: 12.8552 kg''', WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg '''So here we are. 142kg to carry up to 30km altitude.''' The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten, as well as telemetry equipment. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. Rocket parameters are 15.2 mass ratio (9.0 Delta V, 340 constant I_sp), propellants are LOX and E85 at stoichiometric ratio (no venting, no tank insulation), full aluminium structure. Other parameters are the same than above, and the software used was version 1 (see updates here: [[File:Rocket_mass.c]]). [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> 6f0ea3b47fbd94a0deefb3015d26bc96848e067f 436 435 2012-11-28T03:56:35Z Vincent 1 /* Delta V requirements */ gravity drag evaluation wikitext text/x-wiki =First approximations for our rocket= This page summarizes the important numbers about rockets launched from balloon or from aircraft, mainly what minimum vehicle mass we can expect. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following secitons. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: * '''rocket launch altitude''', the higher it is, the less delta V is required to face atmospheric drag and climb against gravity (gravity drag) and actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. ** [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). It still has to be calculated in our conditions. If we need to pitch up after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ** [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch, and it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system has its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph), but its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. The gravity drag is reduced when the thrust is higher because the longer the rocket is in powered flight, the longer it has to overcome the gravity acceleration, as explained [https://en.wikipedia.org/wiki/Gravity_drag here]. Consequently, a higher overall thrust-to-weight ratio will result in less gravity drag. The fact that the engine is heavier in order to provide the higher thrust is still better than having a larger Delta-V to achieve and carrying more propellant (bad mass ratio - see the [[#Effects_of_parameter_changes|graphs]] below). Additionally, the gravity drag [[#Rocket_engine|should]] be partly compensated by small wings on the rocket's fuselage, like on the Pegasus. Since we are in an SSTO perspective, we have to assess how bad the additional mass of these wings is, and also if they can be deployed and jettisoned easily.<br />From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. * '''rocket launch velocity''', which is close to the velocity of the aircraft at rocket release. If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''launch latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, but do we assume that the air is moving at the same speed than the ground? We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. * '''final velocity of the rocket''', which is probably the same than the satellite velocity, except if we decide to catapult it from the rocket somehow, it is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For choice 1: 7800 (orbital speed) + 70 (atmospheric loss) + 700 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8000m/s * For choice 2: 7800 (orbital speed) + 100 (atmospheric loss) + 550 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7750m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. The '''overall thrust-to-weight ratio''' (for the vehicle) will also be required, as it will determine what will be the initial acceleration and the trajectory. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. [[File:Rocket_ignition.png|right]] The rocket release velocity will be the same than the aircraft's, probably meaning a pitch of zero and the rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic, because it would apply very large stress on the vehicle. Moreover, even if it is enough to compensate gravity, it is not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (which not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual, since ground- or balloon-launched rockets start by gaining altitude before orbital speed. These simple examples show how important it is to use the atmosphere to climb. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * '''DRY MASS: 12.8552 kg''', WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg '''So here we are. 142kg to carry up to 30km altitude.''' The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten, as well as telemetry equipment. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. Rocket parameters are 15.2 mass ratio (9.0 Delta V, 340 constant I_sp), propellants are LOX and E85 at stoichiometric ratio (no venting, no tank insulation), full aluminium structure. Other parameters are the same than above, and the software used was version 1 (see updates here: [[File:Rocket_mass.c]]). [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> 52fb58885c5c94d2771d107b9c83fa5d12642a12 442 436 2012-12-11T03:15:38Z Vincent 1 Refactoring the page wikitext text/x-wiki =First approximations for our rocket= This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is obviously to minimize the vehicle mass. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following sections. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * '''DRY MASS: 12.8552 kg''', WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg '''So here we are. 142kg to carry up to 30km altitude.''' The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten, as well as telemetry equipment. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are absolutely not the same, however linear in both cases. Rocket parameters are 15.2 mass ratio (9.0 Delta V, 340 constant I_sp), propellants are LOX and E85 at stoichiometric ratio (no venting, no tank insulation), full aluminium structure. Other parameters are the same than above, and the software used was version 1 (see updates here: [[File:Rocket_mass.c]]). [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These first graphs represent the effect of the payload mass. They use the same data, but the second graph is zoomed-in for very small payloads. For higher values, the graph seems to be linear. At these small values, we can see that there is a slight non-linearity of the vehicle dry and wet masses. The fairing mass was not estimated in our model based on the payload mass, so in reality this is a bit worse. This graph can also apply to any fixed-mass subsystem, like the avionics controller. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The first graph above represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the Isp (propellant efficiency) and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> 475b8216bce853a75d0283ec9b49549afdd0f902 6 77 433 2012-11-27T03:10:25Z Vincent 1 Graph of the rocket trajectory for first moments after release, assuming ignition at the time of the release, a thrust vector with 45 deg pitch, and an overall thrust-to-weight ratio of 2. No atmosphere considered whatsoever. wikitext text/x-wiki Graph of the rocket trajectory for first moments after release, assuming ignition at the time of the release, a thrust vector with 45 deg pitch, and an overall thrust-to-weight ratio of 2. No atmosphere considered whatsoever. 910ca2b783f798fcc1e350bfc9b42813d12aed24 0 17 437 348 2012-11-28T04:12:34Z Vincent 1 /* Cooling for a LOX/E85 engine */ AA LOX film cooling wikitext text/x-wiki This page is an introduction to rocket engines, explaining and comparing some of its subsystems. Our rocket will be based on regular rocket technology, as explained on this page. However some choices need to be made in order to gauge the feasibility of the project as a whole, in particular to have ideas of the possible dimensions of the rocket. These choices are presented on the page of the [[Rocket:First approximations|first approximations for the rocket]]. =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, instead of fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine failure. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotary Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Using LOX for film cooling has also been [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=369#EngineDevelopment demonstrated], by Armadillo Aerospace. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. From SpaceX's experience, it seems that regenerative cooling is sufficient to sustain the combustion temperature hitting the walls. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. feb2bcf15a3cbfc3a2e011b37b6ce02d4da5c554 6 78 441 2012-12-11T00:26:45Z Vincent 1 Rocket model illustration for the computed rocket model: [[File:Rocket_mass.c]] wikitext text/x-wiki Rocket model illustration for the computed rocket model: [[File:Rocket_mass.c]] db21ed85dd16de2565b101cc91205cab83cc862a 6 69 443 388 2012-12-11T03:36:24Z Vincent 1 Adding the picture and few details wikitext text/x-wiki ===Description=== This program is an implementation of the iteration required to compute a rocket's mass. It supports single stage rockets (SSTO) with one engine, takes many input parameters such as thrust-to-weight ratios, propellants data, mass ratio or Isp and Delta V, and computes the dry and wet mass of the rocket. Each subsystem's mass is evaluated and the formulas are simple enough to be changed to suit your needs. [[File:Rocket_model.png|700px]] ===Features=== See the beginning of the file for all input parameters. Each subsystem has a function computing its mass given the rocket length or other parameters, they can be easily modified. See [[Rocket:First_approximations#Our_first_approximation|first approximation]] for the list of output parameters and an example. The mass ratio can be either given or computed from I_sp and Delta V. The formula used is very simple but assumes the exhaust velocity (or I_sp) to be constant [https://en.wikipedia.org/wiki/Rocket_equation#Examples (source)]: e^(-DeltaV/V_e), V_e being the gas exhaust velocity, which can be calculated from Isp: V_e = Isp * g₀. [[File:Rocket_mass_tanks_v1-2.png|right|300px]] The dimensions of the tanks is a major issue. In version 1 and 2, both tanks are computed with the same inner diameter. The goal is to have them roughly the same size so that the rocket can be a simple cylinder. In version 2, tank insulation has been added for the mass concern only. [[File:Rocket_mass_tanks_v3.png|right|300px]] In future version (coming soon) the insulation and wall thickness will be taken into account in the ''outer'' diameter of tanks, so that they can really fit in the same cylinder in any case. Insulation is Aerogel, 13kg/m³ density. The insulation thickness can be calculated from heat and thermal transmission data using this other program: [[File:Heat_transfer_to_rocket_tank.c]]. Tank wall thickness is not properly computed from the pressure and heat, but only estimated as 1mm per 10 atmospheres of pressure, with a minimum value of 1mm for pressurized tanks. The storable fuel tank is currently made of plastic to reduce weight, since pumps will probably be used instead of pressurized tanks. Structural parts are made of aluminium, except the ullage skirt (starting from version 2) which is a carbon/epoxy composite. The vehicle's frame is composed of 5 profiled tubes, 2mm thick, 15mm wide. This ''may'' be enough to support the aircraft staging mechanism and the rocket's weight and flight stress. ===Compilation=== Compile with <code>gcc -O2 -Wall rocket_mass.c -o rocket_mass -lm</code> for normal use, and add <code>-DMAKE_GRAPH</code> after gcc on the command line if you want to create graph data as visible on the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations page]]. bdef29393b38c65f440e66d9c4a9e825aa51668c 444 443 2012-12-11T13:26:57Z Vincent 1 uploaded a new version of "[[File:Rocket mass.c]]":&#32;Refining the model: better structure, more realistic fairing, adding forgotten pipes and valves, tanks alignment on outer shell, composite skirts for non-structural or soft tanks. wikitext text/x-wiki ===Description=== This program is an implementation of the iteration required to compute a rocket's mass. It supports single stage rockets (SSTO) with one engine, takes many input parameters such as thrust-to-weight ratios, propellants data, mass ratio or Isp and Delta V, and computes the dry and wet mass of the rocket. Each subsystem's mass is evaluated and the formulas are simple enough to be changed to suit your needs. [[File:Rocket_model.png|700px]] ===Features=== See the beginning of the file for all input parameters. Each subsystem has a function computing its mass given the rocket length or other parameters, they can be easily modified. See [[Rocket:First_approximations#Our_first_approximation|first approximation]] for the list of output parameters and an example. The mass ratio can be either given or computed from I_sp and Delta V. The formula used is very simple but assumes the exhaust velocity (or I_sp) to be constant [https://en.wikipedia.org/wiki/Rocket_equation#Examples (source)]: e^(-DeltaV/V_e), V_e being the gas exhaust velocity, which can be calculated from Isp: V_e = Isp * g₀. [[File:Rocket_mass_tanks_v1-2.png|right|300px]] The dimensions of the tanks is a major issue. In version 1 and 2, both tanks are computed with the same inner diameter. The goal is to have them roughly the same size so that the rocket can be a simple cylinder. In version 2, tank insulation has been added for the mass concern only. [[File:Rocket_mass_tanks_v3.png|right|300px]] In future version (coming soon) the insulation and wall thickness will be taken into account in the ''outer'' diameter of tanks, so that they can really fit in the same cylinder in any case. Insulation is Aerogel, 13kg/m³ density. The insulation thickness can be calculated from heat and thermal transmission data using this other program: [[File:Heat_transfer_to_rocket_tank.c]]. Tank wall thickness is not properly computed from the pressure and heat, but only estimated as 1mm per 10 atmospheres of pressure, with a minimum value of 1mm for pressurized tanks. The storable fuel tank is currently made of plastic to reduce weight, since pumps will probably be used instead of pressurized tanks. Structural parts are made of aluminium, except the ullage skirt (starting from version 2) which is a carbon/epoxy composite. The vehicle's frame is composed of 5 profiled tubes, 2mm thick, 15mm wide. This ''may'' be enough to support the aircraft staging mechanism and the rocket's weight and flight stress. ===Compilation=== Compile with <code>gcc -O2 -Wall rocket_mass.c -o rocket_mass -lm</code> for normal use, and add <code>-DMAKE_GRAPH</code> after gcc on the command line if you want to create graph data as visible on the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations page]]. bdef29393b38c65f440e66d9c4a9e825aa51668c 445 444 2012-12-11T13:32:57Z Vincent 1 v3 update wikitext text/x-wiki ===Description=== This program is an implementation of the iteration required to compute a rocket's mass. It supports single stage rockets (SSTO) with one engine, takes many input parameters such as thrust-to-weight ratios, propellants data, mass ratio or Isp and Delta V, and computes the dry and wet mass of the rocket. Each subsystem's mass is evaluated and the formulas are simple enough to be changed to suit your needs. The rocket is modelled as depicted below: [[File:Rocket_model.png|700px]] The fairing jettison is currently (v3) not accounted for in the delta V calculation. It is made of aluminium, so it's generally quite heavy, even if it's thin. The pipes and valves have been added in v3 too, also adding an important mass that was forgotten in previous versions. This result in the rocket being around 150kg heavier, because each kg of rocket has approximately 50kg of propellant needed to lift it. The structure's mass has been improved too, taking into account solid tanks for the structure. ===Features=== See the beginning of the file for all input parameters. Each subsystem has a function computing its mass given the rocket length or other parameters, they can be easily modified. See [[Rocket:First_approximations#Our_first_approximation|first approximation]] for the list of output parameters and an example. The mass ratio can be either given or computed from I_sp and Delta V. The formula used is very simple but assumes the exhaust velocity (or I_sp) to be constant [https://en.wikipedia.org/wiki/Rocket_equation#Examples (source)]: e^(-DeltaV/V_e), V_e being the gas exhaust velocity, which can be calculated from Isp: V_e = Isp * g₀. [[File:Rocket_mass_tanks_v1-2.png|right|300px]] The dimensions of the tanks is a major issue. In version 1 and 2, both tanks are computed with the same inner diameter. The goal is to have them roughly the same size so that the rocket can be a simple cylinder. In version 2, tank insulation has been added for the mass concern only. [[File:Rocket_mass_tanks_v3.png|right|300px]] In future version (coming soon) the insulation and wall thickness will be taken into account in the ''outer'' diameter of tanks, so that they can really fit in the same cylinder in any case. Insulation is Aerogel, 13kg/m³ density. The insulation thickness can be calculated from heat and thermal transmission data using this other program: [[File:Heat_transfer_to_rocket_tank.c]]. Tank wall thickness is not properly computed from the pressure and heat, but only estimated as 1mm per 10 atmospheres of pressure, with a minimum value of 1mm for pressurized tanks. The storable fuel tank is currently made of plastic to reduce weight, since pumps will probably be used instead of pressurized tanks. Structural parts are made of aluminium, except the ullage (v2) and soft tank skirts (v3) which is a carbon/epoxy composite. The vehicle's frame is composed of 5 profiled rods, 2mm thick, 15mm wide. This ''may'' be enough to support the aircraft staging mechanism and the rocket's weight and flight stress. ===Compilation=== Compile with <code>gcc -O2 -Wall rocket_mass.c -o rocket_mass -lm</code> for normal use, and add <code>-DMAKE_GRAPH</code> after gcc on the command line if you want to create graph data as visible on the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations page]]. 1746c486ff0a4aa003f76fad82a9f8262759aa76 6 68 446 355 2012-12-11T19:57:44Z Vincent 1 uploaded a new version of "[[File:Payload 0-6k.png]]":&#32;Updated with version 3 of [[File:Rocket_mass.c]]. Delta V is 8300 instead of 9000 m/s too. wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the payload mass (or any other fixed mass subsystem), ranging from 0 to 6 kg payload. 28ef2f319b623642ed4a320b1b90a914668eade3 6 63 447 349 2012-12-11T19:58:48Z Vincent 1 uploaded a new version of "[[File:Payload 0-80kg.png]]":&#32;Updated with version 3 of [[File:Rocket_mass.c]]. Delta V is 8300 instead of 9000 m/s too. wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the payload mass (or any other fixed mass subsystem). 92472815e8c754715a7c221ae990a448ba69138f 448 447 2012-12-11T20:02:24Z Vincent 1 uploaded a new version of "[[File:Payload 0-80kg.png]]":&#32;fixing update wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the payload mass (or any other fixed mass subsystem). 92472815e8c754715a7c221ae990a448ba69138f 6 67 449 354 2012-12-11T20:04:10Z Vincent 1 uploaded a new version of "[[File:Final acceleration.png]]":&#32;Updated with version 3 of [[File:Rocket_mass.c]]. Delta V is 8300 instead of 9000 m/s too. wikitext text/x-wiki Graph representing final acceleration and engine's thrust as function of the initial acceleration (overall thrust-to-weight ratio). c3baab58bab6826bf55de9540823e72d2cd4b055 6 66 450 352 2012-12-11T20:06:18Z Vincent 1 uploaded a new version of "[[File:Overall thrust to weight.png]]":&#32;Updated with version 3 of [[File:Rocket_mass.c]]. Delta V is 8300 instead of 9000 m/s too. wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the vehicle (overall) initial thrust-to-weight ratio. There is an equivalence with the acceleration in number of G's. b2d2d826fab4dc95c2a9f449df1c3c2b38c484f8 6 65 451 351 2012-12-11T20:08:49Z Vincent 1 uploaded a new version of "[[File:Engine thrust to weight.png]]":&#32;Updated with version 3 of [[File:Rocket_mass.c]]. Delta V is 8300 instead of 9000 m/s too, and overall thrust-to-weight ratio is 2.5 instead of 1.25. wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the engine's thrust-to-weight ratio. 629900940158a61aadb92c7445d9fa61534134dc 6 64 452 350 2012-12-11T20:09:22Z Vincent 1 uploaded a new version of "[[File:Mass ratio.png]]":&#32;Updated with version 3 of [[File:Rocket_mass.c]]. Delta V is 8300 instead of 9000 m/s too, and overall thrust-to-weight ratio is 2.5 instead of 1.25. wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the mass ratio (fixed by Isp and Delta V). c170bd9fa96930e3f5461fb37c874b24caedefe4 6 79 453 2012-12-11T20:52:19Z Vincent 1 Graph representing rocket dry and wet mass as function of the Isp. This is clearly exponential on the left because a too low Isp means that orbit cannot be reached, the mass ratio gets too high. Done with version 3 of File:Rocket_mass.c. Delta V is 8300 m wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the Isp. This is clearly exponential on the left because a too low Isp means that orbit cannot be reached, the mass ratio gets too high. Done with version 3 of File:Rocket_mass.c. Delta V is 8300 m/s, overall thrust-to-weight ratio is 2.5. 992b37d988e0a52cfe93e8d96d0db1d4e38dc08b 455 453 2012-12-11T20:54:53Z Vincent 1 fix link wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the Isp. This is clearly exponential on the left because a too low Isp means that orbit cannot be reached, the mass ratio gets too high. Done with version 3 of [[File:Rocket_mass.c]]. Delta V is 8300 m/s, overall thrust-to-weight ratio is 2.5. 3ded20e1f8fd92a818a27695384d890015b3ca93 456 455 2012-12-11T20:59:47Z Vincent 1 uploaded a new version of "[[File:Isp.png]]":&#32;starting at 280 instead of 270 wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the Isp. This is clearly exponential on the left because a too low Isp means that orbit cannot be reached, the mass ratio gets too high. Done with version 3 of [[File:Rocket_mass.c]]. Delta V is 8300 m/s, overall thrust-to-weight ratio is 2.5. 3ded20e1f8fd92a818a27695384d890015b3ca93 6 80 454 2012-12-11T20:53:59Z Vincent 1 Graph representing rocket dry and wet mass as function of the Delta V. Done with version 3 of [[File:Rocket_mass.c]], overall thrust-to-weight ratio is 2.5 and Isp is 340s. wikitext text/x-wiki Graph representing rocket dry and wet mass as function of the Delta V. Done with version 3 of [[File:Rocket_mass.c]], overall thrust-to-weight ratio is 2.5 and Isp is 340s. 56805fbe3018f68126df7b2b779097a93dc35b84 0 61 457 442 2012-12-11T21:06:06Z Vincent 1 /* Effects of parameter changes */ graphs update wikitext text/x-wiki =First approximations for our rocket= This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is obviously to minimize the vehicle mass. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following sections. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * '''DRY MASS: 12.8552 kg''', WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg '''So here we are. 142kg to carry up to 30km altitude.''' The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten, as well as telemetry equipment. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are not the same, although linear in both cases. Rocket parameters are 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. Other parameters are the same as above, and the software used was version 3 (see updates here: [[File:Rocket_mass.c]]). Graphs are valid only for one parameter change, and the evolution of the parameter will be different is other parameters are not the same as here. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet masses. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg'''. [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply impossible to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> 61a634df2a05c48ac00171432362987411a918d9 458 457 2012-12-11T23:14:16Z Vincent 1 /* Our first approximation */ warning message wikitext text/x-wiki =First approximations for our rocket= This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is obviously to minimize the vehicle mass. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following sections. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''''The data in this section is outdated, and was made using the version 1 of the rocket mass program.''''' Results using current version are way worse (around 400kg rocket), but we are waiting for a realistic delta V before changing here. '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * DRY MASS: 12.8552 kg, WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg So here we are. 142kg to carry up to 30km altitude. The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten, as well as telemetry equipment. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are not the same, although linear in both cases. Rocket parameters are 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. Other parameters are the same as above, and the software used was version 3 (see updates here: [[File:Rocket_mass.c]]). Graphs are valid only for one parameter change, and the evolution of the parameter will be different is other parameters are not the same as here. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet masses. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg'''. [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply impossible to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ==References== <references /> 6b7964f69b29c5d4bf2c2dec8047e15b02924e6c 462 458 2012-12-12T04:08:30Z Vincent 1 /* Effects of parameter changes */ the answer to the question of gravity drag and t/w ratio wikitext text/x-wiki =First approximations for our rocket= This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is obviously to minimize the vehicle mass. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following sections. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''''The data in this section is outdated, and was made using the version 1 of the rocket mass program.''''' Results using current version are way worse (around 400kg rocket), but we are waiting for a realistic delta V before changing here. '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * DRY MASS: 12.8552 kg, WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg So here we are. 142kg to carry up to 30km altitude. The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten, as well as telemetry equipment. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are not the same, although linear in both cases. Rocket parameters are 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. Other parameters are the same as above, and the software used was version 3 (see updates here: [[File:Rocket_mass.c]]). Graphs are valid only for one parameter change, and the evolution of the parameter will be different is other parameters are not the same as here. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet masses. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg'''. [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply impossible to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ====The gravity drag and overall thrust-to-weight ratio issue==== The question was: is it better to have a more powerful and heavier engine allowing the gravity drag to be low because the burn time is decreased or to have a lightweight engine burning longer? The answer is in the graph below: it's a compromise. [[File:Gravity_drag_and_thrust_to_weight.png|center]] This graph combines two changing parameters: the overall T/W ratio and the gravity drag that it implies. Changing the gravity drag changes the required delta V, thus changing the mass ratio, that's why the curves have a different minimum. The important for us here is the wet mass, used to design the carrier aircraft. The gravity drag is pretty high compared to what was expected, between 1000 and 1300 m/s in the minimum mass zone. Consequently, we see that the rocket's mass doesn't goes down below 500kg. '''Fixing the overall T/W ratio to 2.5 gives the following rocket: 569.8 kg on the ground''', 46.1 kg dry, 4 kg fairing, engine thrust: 13.93 kN. A two-stage rocket should be considered to lower the mass and size of the rocket (3.53 x 0.47 m tanks here) to something feasible. Some questions follow: is the engine's T/W ratio constant over the thrust range? how does the structure needs to evolve to cope with the stress of the increasing thrust? how is that correct when considering aerodynamics (drag and lift)? ==References== <references /> 10d71ee63472ca7de8eb305a4ff3fbdb709ac5c7 464 462 2012-12-12T18:06:39Z Vincent 1 /* The gravity drag and overall thrust-to-weight ratio issue */ small fixes wikitext text/x-wiki =First approximations for our rocket= This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is obviously to minimize the vehicle mass. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following sections. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''''The data in this section is outdated, and was made using the version 1 of the rocket mass program.''''' Results using current version are way worse (around 400kg rocket), but we are waiting for a realistic delta V before changing here. '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * DRY MASS: 12.8552 kg, WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg So here we are. 142kg to carry up to 30km altitude. The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten, as well as telemetry equipment. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are not the same, although linear in both cases. Rocket parameters are 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. Other parameters are the same as above, and the software used was version 3 (see updates here: [[File:Rocket_mass.c]]). Graphs are valid only for one parameter change, and the evolution of the parameter will be different is other parameters are not the same as here. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet masses. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg'''. [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply impossible to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ====The gravity drag and overall thrust-to-weight ratio issue==== The question was: is it better to have a more powerful and heavier engine allowing the gravity drag to be low because the burn time is decreased or to have a lightweight engine burning longer? The answer is in the graph below: it's a compromise. [[File:Gravity_drag_and_thrust_to_weight.png|center]] This graph combines two changing parameters: the overall T/W ratio and the gravity drag that it implies. Changing the gravity drag changes the required delta V, thus changing the mass ratio, that's why the curves have a different minimum. The important for us here is the wet mass, used to design the carrier aircraft. The gravity drag is pretty high compared to what was expected, between 1000 and 1300 m/s in the minimum mass zone (with Isp = 340s). Consequently, we see that the rocket's mass doesn't go below 550kg. '''Fixing the overall T/W ratio to 2.5 gives the following rocket: 569.8 kg on the ground''', 46.1 kg dry, 4 kg fairing, engine thrust: 13.93 kN. A two-stage rocket should be considered to lower the mass and size of the rocket (3.53 x 0.47 m tanks here) to something feasible. Some questions follow: is the engine's T/W ratio constant over the thrust range? how does the structure needs to evolve to cope with the stress of the increasing thrust? how is that correct when considering aerodynamics (drag and lift)? How does the minimum moves when the Isp changes? ==References== <references /> f73decb0c2e35ec5116e4aa4541df2795a3c599f 465 464 2012-12-12T23:03:00Z Vincent 1 /* Gravity drag */ updated link for selenianboondocks wikitext text/x-wiki =First approximations for our rocket= This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is obviously to minimize the vehicle mass. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following sections. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/ Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== '''''The data in this section is outdated, and was made using the version 1 of the rocket mass program.''''' Results using current version are way worse (around 400kg rocket), but we are waiting for a realistic delta V before changing here. '''Input parameters:''' * Mass ratio: 11.0158, from constant Isp=340s and DeltaV=8000m/s (propellants are 90.92% of the rocket's mass)'' * Length-to-diameter ratio (tanks): 10 * Engine thrust-to-weight ratio: 100. * Overall thrust-to-weight ratio: 1.25 * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics fixed mass: 0.06 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 5mm thick PU foam insulation, 2% venting * Fuel: E85, with 0.5mm plastic tanks (pumping, no pressurization) * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * DRY MASS: 12.8552 kg, WET mass: 141.61 kg * Tanks size (length is rocket length excluding engine and fairing, diameter is the internal tank diameter, nearly rocket's diameter): 2.61112 x 0.261112 m * Engine thrust: 1735.9 N, engine mass: 1.77013 kg * Final thrust-to-weight-ratio: 13.7698 G * Actuators mass: 0.289317 kg, wiring mass: 0.182778 kg, structure mass: 2.70249 kg * LOX tank info: ** diameter: 0.261112 m, length: 1.59797 m ** thickness: 0.002 m, volume: 0.0798638 m^3, mass: 7.23578 kg * E85 tank info: ** diameter: 0.261112 m, length: 1.01314 m ** thickness: 0.0005 m, volume: 0.0506351 m^3, mass: 0.394768 kg So here we are. 142kg to carry up to 30km altitude. The LOX venting has not been properly calculated based on insulation and real vaporization heat data. Pipes and valves for propellant flow have been forgotten, as well as telemetry equipment. The frost created by cryogenic fluids is not taken into account either, and changes the mass ratio. We could increase payload to 1kg to allow for more complex missions than the N-prize mission, and also face unforeseen mass issue (another margin). A multi-stage rocket would be lighter, but more complex to build. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are not the same, although linear in both cases. Rocket parameters are 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. Other parameters are the same as above, and the software used was version 3 (see updates here: [[File:Rocket_mass.c]]). Graphs are valid only for one parameter change, and the evolution of the parameter will be different is other parameters are not the same as here. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet masses. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg'''. [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply impossible to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ====The gravity drag and overall thrust-to-weight ratio issue==== The question was: is it better to have a more powerful and heavier engine allowing the gravity drag to be low because the burn time is decreased or to have a lightweight engine burning longer? The answer is in the graph below: it's a compromise. [[File:Gravity_drag_and_thrust_to_weight.png|center]] This graph combines two changing parameters: the overall T/W ratio and the gravity drag that it implies. Changing the gravity drag changes the required delta V, thus changing the mass ratio, that's why the curves have a different minimum. The important for us here is the wet mass, used to design the carrier aircraft. The gravity drag is pretty high compared to what was expected, between 1000 and 1300 m/s in the minimum mass zone (with Isp = 340s). Consequently, we see that the rocket's mass doesn't go below 550kg. '''Fixing the overall T/W ratio to 2.5 gives the following rocket: 569.8 kg on the ground''', 46.1 kg dry, 4 kg fairing, engine thrust: 13.93 kN. A two-stage rocket should be considered to lower the mass and size of the rocket (3.53 x 0.47 m tanks here) to something feasible. Some questions follow: is the engine's T/W ratio constant over the thrust range? how does the structure needs to evolve to cope with the stress of the increasing thrust? how is that correct when considering aerodynamics (drag and lift)? How does the minimum moves when the Isp changes? ==References== <references /> 0492d4da1e36443367c05adb0e0214d9dc1e9021 466 465 2012-12-13T01:32:42Z Vincent 1 /* Our first approximation */ updating the example with graph's values wikitext text/x-wiki =First approximations for our rocket= This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is obviously to minimize the vehicle mass. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following sections. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/ Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== This approximation has been done with the program in version 3. The input values, delta V of 8.3km/s and T/W ratio of 2.5 are the result of [[#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|the analysis]] of the rocket flight with no atmosphere effect on it (drag or lift). We are working on an aerodynamic analysis that may give more realistic values. There are also the values used for the graphs below. '''Input parameters:''' * Mass ratio: 12.0529, from constant Isp=340s and '''DeltaV=8300m/s''' (propellants are 91.7% of the rocket's mass) * Length-to-diameter ratio (tanks): 8 * Engine thrust-to-weight ratio: 100 * Overall thrust-to-weight ratio: '''2.5''' * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics (computer and battery) fixed mass: 0.25 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 10mm thick Aerogel insulation, 1.25kg venting during ascent * Fuel: E85, with 1mm plastic tanks (pumping, no pressurization) and 1mm composite skirt for outer skin * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * DRY MASS: 43.3571 kg, '''WET mass: 522.579 kg''', fairing mass: 3.828 kg * Tanks external size: 3.42936 x 0.461743 m * Engine thrust: 12811.9 N, engine mass: 13.0645 kg * Final acceleration: '''33.0506 G''' * Actuators mass: 2.45561 kg, wiring mass: 0.240055 kg * Structure mass: 3.42766 kg, valves and pipes: 2.26506 kg * LOX tank info: ** diameter: 0.461743 m, length: 2.12526 m ** wall thickness: 0.002 m, volume: 0.294274 m^3, mass: 15.9799 kg * E85 tank info: ** diameter: 0.461743 m, length: 1.3041 m ** wall thickness: 0.001 m, volume: 0.188839 m^3, mass: 1.77612 kg The frost created by cryogenic fluids is not taken into account. The telemetry equipment has not been seriously evaluated yet. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the scales of the dry (red) and wet (green) curves, they are not the same, although linear in both cases. Rocket parameters are 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. Other parameters are the same as above, and the software used was version 3 (see updates here: [[File:Rocket_mass.c]]). Graphs are valid only for one parameter change, and the evolution of the parameter will be different is other parameters are not the same as here. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet masses. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg'''. [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply impossible to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the rocket's trajectory and release velocity. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition). In our air-to-orbit scenarios, fortunately, we benefit from the aircraft release speed, and the ratio can be lower than 1.3. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the whole engine and thus a mass much larger than the payload's, when the propellants have been consumed the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at least quadratic. ====The gravity drag and overall thrust-to-weight ratio issue==== The question was: is it better to have a more powerful and heavier engine allowing the gravity drag to be low because the burn time is decreased or to have a lightweight engine burning longer? The answer is in the graph below: it's a compromise. [[File:Gravity_drag_and_thrust_to_weight.png|center]] This graph combines two changing parameters: the overall T/W ratio and the gravity drag that it implies. Changing the gravity drag changes the required delta V, thus changing the mass ratio, that's why the curves have a different minimum. The important for us here is the wet mass, used to design the carrier aircraft. The gravity drag is pretty high compared to what was expected, between 1000 and 1300 m/s in the minimum mass zone (with Isp = 340s). Consequently, we see that the rocket's mass doesn't go below 550kg. '''Fixing the overall T/W ratio to 2.5 gives the following rocket: 569.8 kg on the ground''', 46.1 kg dry, 4 kg fairing, engine thrust: 13.93 kN. A two-stage rocket should be considered to lower the mass and size of the rocket (3.53 x 0.47 m tanks here) to something feasible. Some questions follow: is the engine's T/W ratio constant over the thrust range? how does the structure needs to evolve to cope with the stress of the increasing thrust? how is that correct when considering aerodynamics (drag and lift)? How does the minimum moves when the Isp changes? ==References== <references /> 4ce6f41ab182fb7d96a089d0781b6072513e257b 467 466 2012-12-13T01:50:23Z Vincent 1 /* Effects of parameter changes */ text fixes wikitext text/x-wiki =First approximations for our rocket= This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is obviously to minimize the vehicle mass. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following sections. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/ Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. The given delta V doesn't take into account gravity and other forces acting on the vehicle. * the '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''thrust''' is related to this exhaust velocity in the following way {{SERVER}}/images/formulas_mirror/thrust_neg.png. It is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, Ion engines provide very high Isp and v_e but they don't provide enough thrust so they can be used to launch something from Earth to space. * the '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** the '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] an very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== This approximation has been done with the program in version 3. The input values, delta V of 8.3km/s and T/W ratio of 2.5 are the result of [[#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|the analysis]] of the rocket flight with no atmosphere effect on it (drag or lift). We are working on an aerodynamic analysis that may give more realistic values. There are also the values used for the graphs below. '''Input parameters:''' * Mass ratio: 12.0529, from constant Isp=340s and '''DeltaV=8300m/s''' (propellants are 91.7% of the rocket's mass) * Length-to-diameter ratio (tanks): 8 * Engine thrust-to-weight ratio: 100 * Overall thrust-to-weight ratio: '''2.5''' * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics (computer and battery) fixed mass: 0.25 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 10mm thick Aerogel insulation, 1.25kg venting during ascent * Fuel: E85, with 1mm plastic tanks (pumping, no pressurization) and 1mm composite skirt for outer skin * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * DRY MASS: 43.3571 kg, '''WET mass: 522.579 kg''', fairing mass: 3.828 kg * Tanks external size: 3.42936 x 0.461743 m * Engine thrust: 12811.9 N, engine mass: 13.0645 kg * Final acceleration: '''33.0506 G''' * Actuators mass: 2.45561 kg, wiring mass: 0.240055 kg * Structure mass: 3.42766 kg, valves and pipes: 2.26506 kg * LOX tank info: ** diameter: 0.461743 m, length: 2.12526 m ** wall thickness: 0.002 m, volume: 0.294274 m^3, mass: 15.9799 kg * E85 tank info: ** diameter: 0.461743 m, length: 1.3041 m ** wall thickness: 0.001 m, volume: 0.188839 m^3, mass: 1.77612 kg The frost created by cryogenic fluids is not taken into account. The telemetry equipment has not been seriously evaluated yet. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the '''different scales''' of the dry (red) and wet (green) curves. Rocket parameters are the same as above: 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), 2.5 T/W ratio, propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. The software used was version 3 (see updates here: [[File:Rocket_mass.c]]). '''Graphs are valid only for one parameter change''' with the others fixed as above. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet mass curves. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg''' (remember that's only valid with the input parameters as above). [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply unrealistic to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the gravity drag, the aerodynamic lift, the rocket's trajectory and release velocity, and thus hard to evaluate properly. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition) and a lower gravity drag, but a heavier engine. In our air-to-orbit scenarios, fortunately, we may benefit from the aircraft release speed and some lift, and the ratio ''may'' be as low as 1.25. To be studied. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the fully loaded vehicle and thus a mass much larger than the payload's, when the propellants have been consumed and the fairing jettisoned, the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 or 30 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at best quadratic. ====The gravity drag and overall thrust-to-weight ratio issue==== The [[#Gravity_drag|question]] was: is it better to have a more powerful and heavier engine allowing the gravity drag to be low because the burn time is decreased or to have a lightweight engine burning longer? The answer is in the graph below: it's a compromise. [[File:Gravity_drag_and_thrust_to_weight.png|center]] This graph combines two changing parameters: the overall T/W ratio and the gravity drag that it implies. Changing the gravity drag changes the required delta V, thus changing the mass ratio, that's why the curves have a different minimum. The important for us here is the wet mass, used to design the carrier aircraft. The gravity drag is pretty high compared to what was expected, between 1000 and 1300 m/s in the minimum mass zone (with Isp = 340s). Consequently, we see that the rocket's mass doesn't go below 550 kg. Fixing the overall T/W ratio to 2.5 in our model gives the following rocket: 569.8 kg on the ground, 46.1 kg dry, an engine's thrust of 13.93 kN. Other parameters are [[#Our_first_approximation|above]]. A two-stage rocket should be considered to lower the mass and size of the rocket (3.53 x 0.47 m tanks here) to something more feasible. '''Some questions follow:''' is the engine's T/W ratio constant over the thrust range? How does the structure needs to evolve to cope with the stress of the increasing thrust? How is that correct when considering aerodynamics (drag and lift)? How does the minimum moves when the Isp changes? ==References== <references /> cc3b0b09bfa5d3dde30a35f4de4ec1bd0b90b682 470 467 2012-12-20T02:12:53Z Vincent 1 /* Delta V achievement */ more about thrust wikitext text/x-wiki =First approximations for our rocket= This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is obviously to minimize the vehicle mass. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following sections. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999.</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/ Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. Below is a list of important parameters for this equation or for the rocket itself, and how they relate to the delta V. * The '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''Thrust''' is related to this exhaust velocity in the following way: {{SERVER}}/images/formulas_mirror/thrust_neg.png. The dotted ''m'' is the mass flow rate of the engine, i.e. how much propellant is burned every second, and the second part can be ignored for approximations. In vacuum, as the external pressure ''p₀'' gets to 0, the larger the nozzle discharge area is, the higher the thrust, but this has a mass [https://en.wikipedia.org/wiki/Rocket_engine_nozzle#Vacuum_use]. Thrust is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, [https://en.wikipedia.org/wiki/Ion_thruster ion engines] provide very high Isp and exhaust velocity but their mass flow rate is very low, so is thrust. Therefore, they cannot be used to launch something from Earth to space. * The '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** The '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] a very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== This approximation has been done with the program in version 3. The input values, delta V of 8.3km/s and T/W ratio of 2.5 are the result of [[#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|the analysis]] of the rocket flight with no atmosphere effect on it (drag or lift). We are working on an aerodynamic analysis that may give more realistic values. There are also the values used for the graphs below. '''Input parameters:''' * Mass ratio: 12.0529, from constant Isp=340s and '''DeltaV=8300m/s''' (propellants are 91.7% of the rocket's mass) * Length-to-diameter ratio (tanks): 8 * Engine thrust-to-weight ratio: 100 * Overall thrust-to-weight ratio: '''2.5''' * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics (computer and battery) fixed mass: 0.25 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 10mm thick Aerogel insulation, 1.25kg venting during ascent * Fuel: E85, with 1mm plastic tanks (pumping, no pressurization) and 1mm composite skirt for outer skin * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * DRY MASS: 43.3571 kg, '''WET mass: 522.579 kg''', fairing mass: 3.828 kg * Tanks external size: 3.42936 x 0.461743 m * Engine thrust: 12811.9 N, engine mass: 13.0645 kg * Final acceleration: '''33.0506 G''' * Actuators mass: 2.45561 kg, wiring mass: 0.240055 kg * Structure mass: 3.42766 kg, valves and pipes: 2.26506 kg * LOX tank info: ** diameter: 0.461743 m, length: 2.12526 m ** wall thickness: 0.002 m, volume: 0.294274 m^3, mass: 15.9799 kg * E85 tank info: ** diameter: 0.461743 m, length: 1.3041 m ** wall thickness: 0.001 m, volume: 0.188839 m^3, mass: 1.77612 kg The frost created by cryogenic fluids is not taken into account. The telemetry equipment has not been seriously evaluated yet. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the '''different scales''' of the dry (red) and wet (green) curves. Rocket parameters are the same as above: 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), 2.5 T/W ratio, propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. The software used was version 3 (see updates here: [[File:Rocket_mass.c]]). '''Graphs are valid only for one parameter change''' with the others fixed as above. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet mass curves. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg''' (remember that's only valid with the input parameters as above). [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply unrealistic to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the gravity drag, the aerodynamic lift, the rocket's trajectory and release velocity, and thus hard to evaluate properly. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition) and a lower gravity drag, but a heavier engine. In our air-to-orbit scenarios, fortunately, we may benefit from the aircraft release speed and some lift, and the ratio ''may'' be as low as 1.25. To be studied. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the fully loaded vehicle and thus a mass much larger than the payload's, when the propellants have been consumed and the fairing jettisoned, the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 or 30 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at best quadratic. ====The gravity drag and overall thrust-to-weight ratio issue==== The [[#Gravity_drag|question]] was: is it better to have a more powerful and heavier engine allowing the gravity drag to be low because the burn time is decreased or to have a lightweight engine burning longer? The answer is in the graph below: it's a compromise. [[File:Gravity_drag_and_thrust_to_weight.png|center]] This graph combines two changing parameters: the overall T/W ratio and the gravity drag that it implies. Changing the gravity drag changes the required delta V, thus changing the mass ratio, that's why the curves have a different minimum. The important for us here is the wet mass, used to design the carrier aircraft. The gravity drag is pretty high compared to what was expected, between 1000 and 1300 m/s in the minimum mass zone (with Isp = 340s). Consequently, we see that the rocket's mass doesn't go below 550 kg. Fixing the overall T/W ratio to 2.5 in our model gives the following rocket: 569.8 kg on the ground, 46.1 kg dry, an engine's thrust of 13.93 kN. Other parameters are [[#Our_first_approximation|above]]. A two-stage rocket should be considered to lower the mass and size of the rocket (3.53 x 0.47 m tanks here) to something more feasible. '''Some questions follow:''' is the engine's T/W ratio constant over the thrust range? How does the structure needs to evolve to cope with the stress of the increasing thrust? How is that correct when considering aerodynamics (drag and lift)? How does the minimum moves when the Isp changes? ==References== <references /> bc2e265a0690b84abf52107dce6570269636943e 471 470 2012-12-20T02:24:11Z Vincent 1 adding "the Thrust and Specific Impulse" reference wikitext text/x-wiki =First approximations for our rocket= This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is obviously to minimize the vehicle mass. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following sections. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999. ([http://dspace.mit.edu/handle/1721.1/9383 download])</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/ Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] tells us that {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where m₀ is the initial total mass including propellant, m₁ is the final total mass (dry rocket mass), and v_e is the effective exhaust velocity. Below is a list of important parameters for this equation or for the rocket itself, and how they relate to the delta V. * The '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png <ref name="thrust_and_isp">Prof. Z. S. Spakovszky. Unified Thermodynamics and Propulsion, [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/node102.html section 14.1] '''Thrust and Specific Impulse for Rockets'''. MIT teachings, December 2007</ref>. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''Thrust''' is related to this exhaust velocity in the following way: {{SERVER}}/images/formulas_mirror/thrust_neg.png <ref name="thrust_and_isp" />. The dotted ''m'' is the mass flow rate of the engine, i.e. how much propellant is burned every second, and the second part can be ignored for approximations. In vacuum, as the external pressure ''p₀'' gets to 0, the larger the nozzle discharge area is, the higher the thrust, but this has a mass [https://en.wikipedia.org/wiki/Rocket_engine_nozzle#Vacuum_use]. Thrust is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, [https://en.wikipedia.org/wiki/Ion_thruster ion engines] provide very high Isp and exhaust velocity but their mass flow rate is very low, so is thrust. Therefore, they cannot be used to launch something from Earth to space. * The '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** The '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] a very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== This approximation has been done with the program in version 3. The input values, delta V of 8.3km/s and T/W ratio of 2.5 are the result of [[#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|the analysis]] of the rocket flight with no atmosphere effect on it (drag or lift). We are working on an aerodynamic analysis that may give more realistic values. There are also the values used for the graphs below. '''Input parameters:''' * Mass ratio: 12.0529, from constant Isp=340s and '''DeltaV=8300m/s''' (propellants are 91.7% of the rocket's mass) * Length-to-diameter ratio (tanks): 8 * Engine thrust-to-weight ratio: 100 * Overall thrust-to-weight ratio: '''2.5''' * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics (computer and battery) fixed mass: 0.25 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 10mm thick Aerogel insulation, 1.25kg venting during ascent * Fuel: E85, with 1mm plastic tanks (pumping, no pressurization) and 1mm composite skirt for outer skin * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * DRY MASS: 43.3571 kg, '''WET mass: 522.579 kg''', fairing mass: 3.828 kg * Tanks external size: 3.42936 x 0.461743 m * Engine thrust: 12811.9 N, engine mass: 13.0645 kg * Final acceleration: '''33.0506 G''' * Actuators mass: 2.45561 kg, wiring mass: 0.240055 kg * Structure mass: 3.42766 kg, valves and pipes: 2.26506 kg * LOX tank info: ** diameter: 0.461743 m, length: 2.12526 m ** wall thickness: 0.002 m, volume: 0.294274 m^3, mass: 15.9799 kg * E85 tank info: ** diameter: 0.461743 m, length: 1.3041 m ** wall thickness: 0.001 m, volume: 0.188839 m^3, mass: 1.77612 kg The frost created by cryogenic fluids is not taken into account. The telemetry equipment has not been seriously evaluated yet. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the '''different scales''' of the dry (red) and wet (green) curves. Rocket parameters are the same as above: 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), 2.5 T/W ratio, propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. The software used was version 3 (see updates here: [[File:Rocket_mass.c]]). '''Graphs are valid only for one parameter change''' with the others fixed as above. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet mass curves. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg''' (remember that's only valid with the input parameters as above). [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply unrealistic to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the gravity drag, the aerodynamic lift, the rocket's trajectory and release velocity, and thus hard to evaluate properly. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition) and a lower gravity drag, but a heavier engine. In our air-to-orbit scenarios, fortunately, we may benefit from the aircraft release speed and some lift, and the ratio ''may'' be as low as 1.25. To be studied. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the fully loaded vehicle and thus a mass much larger than the payload's, when the propellants have been consumed and the fairing jettisoned, the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 or 30 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at best quadratic. ====The gravity drag and overall thrust-to-weight ratio issue==== The [[#Gravity_drag|question]] was: is it better to have a more powerful and heavier engine allowing the gravity drag to be low because the burn time is decreased or to have a lightweight engine burning longer? The answer is in the graph below: it's a compromise. [[File:Gravity_drag_and_thrust_to_weight.png|center]] This graph combines two changing parameters: the overall T/W ratio and the gravity drag that it implies. Changing the gravity drag changes the required delta V, thus changing the mass ratio, that's why the curves have a different minimum. The important for us here is the wet mass, used to design the carrier aircraft. The gravity drag is pretty high compared to what was expected, between 1000 and 1300 m/s in the minimum mass zone (with Isp = 340s). Consequently, we see that the rocket's mass doesn't go below 550 kg. Fixing the overall T/W ratio to 2.5 in our model gives the following rocket: 569.8 kg on the ground, 46.1 kg dry, an engine's thrust of 13.93 kN. Other parameters are [[#Our_first_approximation|above]]. A two-stage rocket should be considered to lower the mass and size of the rocket (3.53 x 0.47 m tanks here) to something more feasible. '''Some questions follow:''' is the engine's T/W ratio constant over the thrust range? How does the structure needs to evolve to cope with the stress of the increasing thrust? How is that correct when considering aerodynamics (drag and lift)? How does the minimum moves when the Isp changes? ==References== <references /> 85e27418bfc36fec41d1bb18b85ab765d3665e7e 472 471 2012-12-20T03:09:11Z Vincent 1 /* Delta V achievement */ simplifed rocket equation issue wikitext text/x-wiki =First approximations for our rocket= This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is obviously to minimize the vehicle mass. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following sections. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999. ([http://dspace.mit.edu/handle/1721.1/9383 download])</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/ Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The ''simplified'' [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] is expressed as {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where ''m₀'' is the initial total mass including propellant (rocket's wet mass), ''m₁'' is the final total mass (rocket's dry mass), and ''v_e'' is the effective exhaust velocity. This simplification is only '''valid for''': no gravity nor aerodynamic drag, constant exhaust velocity (or Isp), ideal nozzle expansion (''p_e'' = ''p₀''), '''and with an initial null velocity''' [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/node103.html]. The calculations and model described below have been done with this simplified equation, yet it is incorrect since in vacuum we have no ideal nozzle expansion and the initial velocity is not null for aircraft-launched rockets. The effect of these two parameters have to be evaluated soon. Below is a list of important parameters for this equation or for the rocket itself, and how they relate to the delta V. * The '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png <ref name="thrust_and_isp">Prof. Z. S. Spakovszky. Unified Thermodynamics and Propulsion, [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/node102.html section 14.1] '''Thrust and Specific Impulse for Rockets'''. MIT teachings, December 2007</ref>. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''Thrust''' is related to this exhaust velocity in the following way: {{SERVER}}/images/formulas_mirror/thrust_neg.png <ref name="thrust_and_isp" />. The dotted ''m'' is the mass flow rate of the engine, i.e. how much propellant is burned every second, and the second part can be ignored for approximations. In vacuum, as the external pressure ''p₀'' gets to 0, the larger the nozzle discharge area is, the higher the thrust, but this has a mass [https://en.wikipedia.org/wiki/Rocket_engine_nozzle#Vacuum_use]. Thrust is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, [https://en.wikipedia.org/wiki/Ion_thruster ion engines] provide very high Isp and exhaust velocity but their mass flow rate is very low, so is thrust. Therefore, they cannot be used to launch something from Earth to space. * The '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** The '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] a very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== This approximation has been done with the program in version 3. The input values, delta V of 8.3km/s and T/W ratio of 2.5 are the result of [[#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|the analysis]] of the rocket flight with no atmosphere effect on it (drag or lift). We are working on an aerodynamic analysis that may give more realistic values. There are also the values used for the graphs below. '''Input parameters:''' * Mass ratio: 12.0529, from constant Isp=340s and '''DeltaV=8300m/s''' (propellants are 91.7% of the rocket's mass) * Length-to-diameter ratio (tanks): 8 * Engine thrust-to-weight ratio: 100 * Overall thrust-to-weight ratio: '''2.5''' * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics (computer and battery) fixed mass: 0.25 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 10mm thick Aerogel insulation, 1.25kg venting during ascent * Fuel: E85, with 1mm plastic tanks (pumping, no pressurization) and 1mm composite skirt for outer skin * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * DRY MASS: 43.3571 kg, '''WET mass: 522.579 kg''', fairing mass: 3.828 kg * Tanks external size: 3.42936 x 0.461743 m * Engine thrust: 12811.9 N, engine mass: 13.0645 kg * Final acceleration: '''33.0506 G''' * Actuators mass: 2.45561 kg, wiring mass: 0.240055 kg * Structure mass: 3.42766 kg, valves and pipes: 2.26506 kg * LOX tank info: ** diameter: 0.461743 m, length: 2.12526 m ** wall thickness: 0.002 m, volume: 0.294274 m^3, mass: 15.9799 kg * E85 tank info: ** diameter: 0.461743 m, length: 1.3041 m ** wall thickness: 0.001 m, volume: 0.188839 m^3, mass: 1.77612 kg The frost created by cryogenic fluids is not taken into account. The telemetry equipment has not been seriously evaluated yet. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the '''different scales''' of the dry (red) and wet (green) curves. Rocket parameters are the same as above: 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), 2.5 T/W ratio, propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. The software used was version 3 (see updates here: [[File:Rocket_mass.c]]). '''Graphs are valid only for one parameter change''' with the others fixed as above. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet mass curves. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg''' (remember that's only valid with the input parameters as above). [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply unrealistic to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the gravity drag, the aerodynamic lift, the rocket's trajectory and release velocity, and thus hard to evaluate properly. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition) and a lower gravity drag, but a heavier engine. In our air-to-orbit scenarios, fortunately, we may benefit from the aircraft release speed and some lift, and the ratio ''may'' be as low as 1.25. To be studied. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the fully loaded vehicle and thus a mass much larger than the payload's, when the propellants have been consumed and the fairing jettisoned, the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 or 30 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at best quadratic. ====The gravity drag and overall thrust-to-weight ratio issue==== The [[#Gravity_drag|question]] was: is it better to have a more powerful and heavier engine allowing the gravity drag to be low because the burn time is decreased or to have a lightweight engine burning longer? The answer is in the graph below: it's a compromise. [[File:Gravity_drag_and_thrust_to_weight.png|center]] This graph combines two changing parameters: the overall T/W ratio and the gravity drag that it implies. Changing the gravity drag changes the required delta V, thus changing the mass ratio, that's why the curves have a different minimum. The important for us here is the wet mass, used to design the carrier aircraft. The gravity drag is pretty high compared to what was expected, between 1000 and 1300 m/s in the minimum mass zone (with Isp = 340s). Consequently, we see that the rocket's mass doesn't go below 550 kg. Fixing the overall T/W ratio to 2.5 in our model gives the following rocket: 569.8 kg on the ground, 46.1 kg dry, an engine's thrust of 13.93 kN. Other parameters are [[#Our_first_approximation|above]]. A two-stage rocket should be considered to lower the mass and size of the rocket (3.53 x 0.47 m tanks here) to something more feasible. '''Some questions follow:''' is the engine's T/W ratio constant over the thrust range? How does the structure needs to evolve to cope with the stress of the increasing thrust? How is that correct when considering aerodynamics (drag and lift)? How does the minimum moves when the Isp changes? ==References== <references /> a48bf7c55dcb621e11d51adeb40552301a786e02 473 472 2012-12-25T00:36:32Z Vincent 1 /* The gravity drag and overall thrust-to-weight ratio issue */ smaller engine is better wikitext text/x-wiki =First approximations for our rocket= This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is obviously to minimize the vehicle mass. The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. More study is required to derive a minimum mass estimation from this, done in following sections. This information is used as a first approximation for the design of other systems, like the aircraft and its engines. Information about particular systems of rocket engines can be found in the [[RocketEngines|rocket engines]] page. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999. ([http://dspace.mit.edu/handle/1721.1/9383 download])</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/ Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The ''simplified'' [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] is expressed as {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where ''m₀'' is the initial total mass including propellant (rocket's wet mass), ''m₁'' is the final total mass (rocket's dry mass), and ''v_e'' is the effective exhaust velocity. This simplification is only '''valid for''': no gravity nor aerodynamic drag, constant exhaust velocity (or Isp), ideal nozzle expansion (''p_e'' = ''p₀''), '''and with an initial null velocity''' [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/node103.html]. The calculations and model described below have been done with this simplified equation, yet it is incorrect since in vacuum we have no ideal nozzle expansion and the initial velocity is not null for aircraft-launched rockets. The effect of these two parameters have to be evaluated soon. Below is a list of important parameters for this equation or for the rocket itself, and how they relate to the delta V. * The '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png <ref name="thrust_and_isp">Prof. Z. S. Spakovszky. Unified Thermodynamics and Propulsion, [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/node102.html section 14.1] '''Thrust and Specific Impulse for Rockets'''. MIT teachings, December 2007</ref>. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''Thrust''' is related to this exhaust velocity in the following way: {{SERVER}}/images/formulas_mirror/thrust_neg.png <ref name="thrust_and_isp" />. The dotted ''m'' is the mass flow rate of the engine, i.e. how much propellant is burned every second, and the second part can be ignored for approximations. In vacuum, as the external pressure ''p₀'' gets to 0, the larger the nozzle discharge area is, the higher the thrust, but this has a mass [https://en.wikipedia.org/wiki/Rocket_engine_nozzle#Vacuum_use]. Thrust is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, [https://en.wikipedia.org/wiki/Ion_thruster ion engines] provide very high Isp and exhaust velocity but their mass flow rate is very low, so is thrust. Therefore, they cannot be used to launch something from Earth to space. * The '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** The '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] a very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== This approximation has been done with the program in version 3. The input values, delta V of 8.3km/s and T/W ratio of 2.5 are the result of [[#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|the analysis]] of the rocket flight with no atmosphere effect on it (drag or lift). We are working on an aerodynamic analysis that may give more realistic values. There are also the values used for the graphs below. '''Input parameters:''' * Mass ratio: 12.0529, from constant Isp=340s and '''DeltaV=8300m/s''' (propellants are 91.7% of the rocket's mass) * Length-to-diameter ratio (tanks): 8 * Engine thrust-to-weight ratio: 100 * Overall thrust-to-weight ratio: '''2.5''' * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics (computer and battery) fixed mass: 0.25 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 10mm thick Aerogel insulation, 1.25kg venting during ascent * Fuel: E85, with 1mm plastic tanks (pumping, no pressurization) and 1mm composite skirt for outer skin * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * DRY MASS: 43.3571 kg, '''WET mass: 522.579 kg''', fairing mass: 3.828 kg * Tanks external size: 3.42936 x 0.461743 m * Engine thrust: 12811.9 N, engine mass: 13.0645 kg * Final acceleration: '''33.0506 G''' * Actuators mass: 2.45561 kg, wiring mass: 0.240055 kg * Structure mass: 3.42766 kg, valves and pipes: 2.26506 kg * LOX tank info: ** diameter: 0.461743 m, length: 2.12526 m ** wall thickness: 0.002 m, volume: 0.294274 m^3, mass: 15.9799 kg * E85 tank info: ** diameter: 0.461743 m, length: 1.3041 m ** wall thickness: 0.001 m, volume: 0.188839 m^3, mass: 1.77612 kg The frost created by cryogenic fluids is not taken into account. The telemetry equipment has not been seriously evaluated yet. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the '''different scales''' of the dry (red) and wet (green) curves. Rocket parameters are the same as above: 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), 2.5 T/W ratio, propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. The software used was version 3 (see updates here: [[File:Rocket_mass.c]]). '''Graphs are valid only for one parameter change''' with the others fixed as above. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet mass curves. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg''' (remember that's only valid with the input parameters as above). [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply unrealistic to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the gravity drag, the aerodynamic lift, the rocket's trajectory and release velocity, and thus hard to evaluate properly. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition) and a lower gravity drag, but a heavier engine. In our air-to-orbit scenarios, fortunately, we may benefit from the aircraft release speed and some lift, and the ratio ''may'' be as low as 1.25. To be studied. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the fully loaded vehicle and thus a mass much larger than the payload's, when the propellants have been consumed and the fairing jettisoned, the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 or 30 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at best quadratic. ====The gravity drag and overall thrust-to-weight ratio issue==== The [[#Gravity_drag|question]] was: ''is it better to have a more powerful and heavier engine that minimizes the gravity drag because the burn time is decreased or to have a lightweight engine burning longer?'' An issue with this question is the term ''better''. It was first thought as the vehicle with the lowest mass and in this case the answer is in the graph below: it's a compromise. However, when cost is considered, a less powerful engine costs less to manufacture and to design, and it should be evaluated how much more fuel can be taken to compensate this smaller engine without increasing the overall cost. That's a complicated question too. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/comment-page-2/#comment-12397 This discussion] confirms the fact that a 1.25 T/W ratio is better costwise. [[File:Gravity_drag_and_thrust_to_weight.png|center]] This graph combines two changing parameters: the overall T/W ratio and the gravity drag that it implies. Changing the gravity drag changes the required delta V, thus changing the mass ratio, that's why the curves have a different minimum. The concerning parameter here is the wet mass, used to design the carrier aircraft. The gravity drag is pretty high compared to what was expected, between 1000 and 1300 m/s in the minimum mass zone (with Isp = 340s). Consequently, we see that the rocket's mass doesn't go below 550 kg. Incidentally, the wet mass doesn't evolve much for T/W ratios between 1.8 and 3.2, and designing a smaller engine because of the cost would not have much bad consequences on the rest of the vehicle. Fixing the overall T/W ratio to 2.5 in our model gives the following rocket: 569.8 kg on the ground, 46.1 kg dry, an engine's thrust of 13.93 kN. Other parameters are [[#Our_first_approximation|above]]. A two-stage rocket should be considered to lower the mass and size of the rocket (3.53 x 0.47 m tanks here) to something more feasible. '''Some questions follow:''' is the engine's T/W ratio constant over the thrust range? How does the structure needs to evolve to cope with the stress of the increasing thrust? How is that correct when considering aerodynamics (drag and lift)? How does the minimum moves when the Isp changes? ==References== <references /> bbad4df8920a40730dfe05b95a3daaedff15da4b 474 473 2012-12-29T00:54:10Z Vincent 1 New introduction wikitext text/x-wiki This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is to minimize the cost of the vehicle and its carrier aircraft. Minimizing the mass is generally the good way to design a rocket, although it may increase the cost when some parameters are changed, like the rocket's thrust. A trade-off has to be made as explained on this page. Information about particular systems of existing rocket engines can be found in the [[RocketEngines|rocket engines]] page. =First mass approximation for our rocket= The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. Combined with a first evaluation of the mass of [[#Systems_design_for_minimum_mass|each rocket system]], a computer model for rocket mass has been [[#Minimum_mass_evaluation|made]]. The effect of the variation of single parameter on the wet and dry mass of a rocket has been documented as graphs using this model. The gravity drag is a particular parameter affected by other parameters and its variation has dramatic consequences on the mass of the vehicle. It is studied more deeply at the [[#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|bottom of the page]]. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999. ([http://dspace.mit.edu/handle/1721.1/9383 download])</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/ Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The ''simplified'' [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] is expressed as {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where ''m₀'' is the initial total mass including propellant (rocket's wet mass), ''m₁'' is the final total mass (rocket's dry mass), and ''v_e'' is the effective exhaust velocity. This simplification is only '''valid for''': no gravity nor aerodynamic drag, constant exhaust velocity (or Isp), ideal nozzle expansion (''p_e'' = ''p₀''), '''and with an initial null velocity''' [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/node103.html]. The calculations and model described below have been done with this simplified equation, yet it is incorrect since in vacuum we have no ideal nozzle expansion and the initial velocity is not null for aircraft-launched rockets. The effect of these two parameters have to be evaluated soon. Below is a list of important parameters for this equation or for the rocket itself, and how they relate to the delta V. * The '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png <ref name="thrust_and_isp">Prof. Z. S. Spakovszky. Unified Thermodynamics and Propulsion, [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/node102.html section 14.1] '''Thrust and Specific Impulse for Rockets'''. MIT teachings, December 2007</ref>. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''Thrust''' is related to this exhaust velocity in the following way: {{SERVER}}/images/formulas_mirror/thrust_neg.png <ref name="thrust_and_isp" />. The dotted ''m'' is the mass flow rate of the engine, i.e. how much propellant is burned every second, and the second part can be ignored for approximations. In vacuum, as the external pressure ''p₀'' gets to 0, the larger the nozzle discharge area is, the higher the thrust, but this has a mass [https://en.wikipedia.org/wiki/Rocket_engine_nozzle#Vacuum_use]. Thrust is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, [https://en.wikipedia.org/wiki/Ion_thruster ion engines] provide very high Isp and exhaust velocity but their mass flow rate is very low, so is thrust. Therefore, they cannot be used to launch something from Earth to space. * The '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** The '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] a very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== This approximation has been done with the program in version 3. The input values, delta V of 8.3km/s and T/W ratio of 2.5 are the result of [[#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|the analysis]] of the rocket flight with no atmosphere effect on it (drag or lift). We are working on an aerodynamic analysis that may give more realistic values. There are also the values used for the graphs below. '''Input parameters:''' * Mass ratio: 12.0529, from constant Isp=340s and '''DeltaV=8300m/s''' (propellants are 91.7% of the rocket's mass) * Length-to-diameter ratio (tanks): 8 * Engine thrust-to-weight ratio: 100 * Overall thrust-to-weight ratio: '''2.5''' * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics (computer and battery) fixed mass: 0.25 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 10mm thick Aerogel insulation, 1.25kg venting during ascent * Fuel: E85, with 1mm plastic tanks (pumping, no pressurization) and 1mm composite skirt for outer skin * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * DRY MASS: 43.3571 kg, '''WET mass: 522.579 kg''', fairing mass: 3.828 kg * Tanks external size: 3.42936 x 0.461743 m * Engine thrust: 12811.9 N, engine mass: 13.0645 kg * Final acceleration: '''33.0506 G''' * Actuators mass: 2.45561 kg, wiring mass: 0.240055 kg * Structure mass: 3.42766 kg, valves and pipes: 2.26506 kg * LOX tank info: ** diameter: 0.461743 m, length: 2.12526 m ** wall thickness: 0.002 m, volume: 0.294274 m^3, mass: 15.9799 kg * E85 tank info: ** diameter: 0.461743 m, length: 1.3041 m ** wall thickness: 0.001 m, volume: 0.188839 m^3, mass: 1.77612 kg The frost created by cryogenic fluids is not taken into account. The telemetry equipment has not been seriously evaluated yet. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the '''different scales''' of the dry (red) and wet (green) curves. Rocket parameters are the same as above: 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), 2.5 T/W ratio, propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. The software used was version 3 (see updates here: [[File:Rocket_mass.c]]). '''Graphs are valid only for one parameter change''' with the others fixed as above. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet mass curves. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg''' (remember that's only valid with the input parameters as above). [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply unrealistic to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the gravity drag, the aerodynamic lift, the rocket's trajectory and release velocity, and thus hard to evaluate properly. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition) and a lower gravity drag, but a heavier engine. In our air-to-orbit scenarios, fortunately, we may benefit from the aircraft release speed and some lift, and the ratio ''may'' be as low as 1.25. To be studied. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the fully loaded vehicle and thus a mass much larger than the payload's, when the propellants have been consumed and the fairing jettisoned, the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 or 30 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at best quadratic. ====The gravity drag and overall thrust-to-weight ratio issue==== The [[#Gravity_drag|question]] was: ''is it better to have a more powerful and heavier engine that minimizes the gravity drag because the burn time is decreased or to have a lightweight engine burning longer?'' An issue with this question is the term ''better''. It was first thought as the vehicle with the lowest mass and in this case the answer is in the graph below: it's a compromise. However, when cost is considered, a less powerful engine costs less to manufacture and to design, and it should be evaluated how much more fuel can be taken to compensate this smaller engine without increasing the overall cost. That's a complicated question too. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/comment-page-2/#comment-12397 This discussion] confirms the fact that a 1.25 T/W ratio is better costwise. [[File:Gravity_drag_and_thrust_to_weight.png|center]] This graph combines two changing parameters: the overall T/W ratio and the gravity drag that it implies. Changing the gravity drag changes the required delta V, thus changing the mass ratio, that's why the curves have a different minimum. The concerning parameter here is the wet mass, used to design the carrier aircraft. The gravity drag is pretty high compared to what was expected, between 1000 and 1300 m/s in the minimum mass zone (with Isp = 340s). Consequently, we see that the rocket's mass doesn't go below 550 kg. Incidentally, the wet mass doesn't evolve much for T/W ratios between 1.8 and 3.2, and designing a smaller engine because of the cost would not have much bad consequences on the rest of the vehicle. Fixing the overall T/W ratio to 2.5 in our model gives the following rocket: 569.8 kg on the ground, 46.1 kg dry, an engine's thrust of 13.93 kN. Other parameters are [[#Our_first_approximation|above]]. A two-stage rocket should be considered to lower the mass and size of the rocket (3.53 x 0.47 m tanks here) to something more feasible. '''Some questions follow:''' is the engine's T/W ratio constant over the thrust range? How does the structure needs to evolve to cope with the stress of the increasing thrust? How is that correct when considering aerodynamics (drag and lift)? How does the minimum moves when the Isp changes? ==References== <references /> 0f72f2997900ce2e281bf6c15cbe5b7becfd62be 475 474 2012-12-29T00:56:11Z Vincent 1 adding rocket category wikitext text/x-wiki This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is to minimize the cost of the vehicle and its carrier aircraft. Minimizing the mass is generally the good way to design a rocket, although it may increase the cost when some parameters are changed, like the rocket's thrust. A trade-off has to be made as explained on this page. Information about particular systems of existing rocket engines can be found in the [[RocketEngines|rocket engines]] page. =First mass approximation for our rocket= The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. Combined with a first evaluation of the mass of [[#Systems_design_for_minimum_mass|each rocket system]], a computer model for rocket mass has been [[#Minimum_mass_evaluation|made]]. The effect of the variation of single parameter on the wet and dry mass of a rocket has been documented as graphs using this model. The gravity drag is a particular parameter affected by other parameters and its variation has dramatic consequences on the mass of the vehicle. It is studied more deeply at the [[#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|bottom of the page]]. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999. ([http://dspace.mit.edu/handle/1721.1/9383 download])</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/ Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The ''simplified'' [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] is expressed as {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where ''m₀'' is the initial total mass including propellant (rocket's wet mass), ''m₁'' is the final total mass (rocket's dry mass), and ''v_e'' is the effective exhaust velocity. This simplification is only '''valid for''': no gravity nor aerodynamic drag, constant exhaust velocity (or Isp), ideal nozzle expansion (''p_e'' = ''p₀''), '''and with an initial null velocity''' [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/node103.html]. The calculations and model described below have been done with this simplified equation, yet it is incorrect since in vacuum we have no ideal nozzle expansion and the initial velocity is not null for aircraft-launched rockets. The effect of these two parameters have to be evaluated soon. Below is a list of important parameters for this equation or for the rocket itself, and how they relate to the delta V. * The '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png <ref name="thrust_and_isp">Prof. Z. S. Spakovszky. Unified Thermodynamics and Propulsion, [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/node102.html section 14.1] '''Thrust and Specific Impulse for Rockets'''. MIT teachings, December 2007</ref>. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''Thrust''' is related to this exhaust velocity in the following way: {{SERVER}}/images/formulas_mirror/thrust_neg.png <ref name="thrust_and_isp" />. The dotted ''m'' is the mass flow rate of the engine, i.e. how much propellant is burned every second, and the second part can be ignored for approximations. In vacuum, as the external pressure ''p₀'' gets to 0, the larger the nozzle discharge area is, the higher the thrust, but this has a mass [https://en.wikipedia.org/wiki/Rocket_engine_nozzle#Vacuum_use]. Thrust is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, [https://en.wikipedia.org/wiki/Ion_thruster ion engines] provide very high Isp and exhaust velocity but their mass flow rate is very low, so is thrust. Therefore, they cannot be used to launch something from Earth to space. * The '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** The '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] a very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== This approximation has been done with the program in version 3. The input values, delta V of 8.3km/s and T/W ratio of 2.5 are the result of [[#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|the analysis]] of the rocket flight with no atmosphere effect on it (drag or lift). We are working on an aerodynamic analysis that may give more realistic values. There are also the values used for the graphs below. '''Input parameters:''' * Mass ratio: 12.0529, from constant Isp=340s and '''DeltaV=8300m/s''' (propellants are 91.7% of the rocket's mass) * Length-to-diameter ratio (tanks): 8 * Engine thrust-to-weight ratio: 100 * Overall thrust-to-weight ratio: '''2.5''' * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics (computer and battery) fixed mass: 0.25 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 10mm thick Aerogel insulation, 1.25kg venting during ascent * Fuel: E85, with 1mm plastic tanks (pumping, no pressurization) and 1mm composite skirt for outer skin * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * DRY MASS: 43.3571 kg, '''WET mass: 522.579 kg''', fairing mass: 3.828 kg * Tanks external size: 3.42936 x 0.461743 m * Engine thrust: 12811.9 N, engine mass: 13.0645 kg * Final acceleration: '''33.0506 G''' * Actuators mass: 2.45561 kg, wiring mass: 0.240055 kg * Structure mass: 3.42766 kg, valves and pipes: 2.26506 kg * LOX tank info: ** diameter: 0.461743 m, length: 2.12526 m ** wall thickness: 0.002 m, volume: 0.294274 m^3, mass: 15.9799 kg * E85 tank info: ** diameter: 0.461743 m, length: 1.3041 m ** wall thickness: 0.001 m, volume: 0.188839 m^3, mass: 1.77612 kg The frost created by cryogenic fluids is not taken into account. The telemetry equipment has not been seriously evaluated yet. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the '''different scales''' of the dry (red) and wet (green) curves. Rocket parameters are the same as above: 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), 2.5 T/W ratio, propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. The software used was version 3 (see updates here: [[File:Rocket_mass.c]]). '''Graphs are valid only for one parameter change''' with the others fixed as above. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet mass curves. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg''' (remember that's only valid with the input parameters as above). [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply unrealistic to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the gravity drag, the aerodynamic lift, the rocket's trajectory and release velocity, and thus hard to evaluate properly. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition) and a lower gravity drag, but a heavier engine. In our air-to-orbit scenarios, fortunately, we may benefit from the aircraft release speed and some lift, and the ratio ''may'' be as low as 1.25. To be studied. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the fully loaded vehicle and thus a mass much larger than the payload's, when the propellants have been consumed and the fairing jettisoned, the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 or 30 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at best quadratic. ====The gravity drag and overall thrust-to-weight ratio issue==== The [[#Gravity_drag|question]] was: ''is it better to have a more powerful and heavier engine that minimizes the gravity drag because the burn time is decreased or to have a lightweight engine burning longer?'' An issue with this question is the term ''better''. It was first thought as the vehicle with the lowest mass and in this case the answer is in the graph below: it's a compromise. However, when cost is considered, a less powerful engine costs less to manufacture and to design, and it should be evaluated how much more fuel can be taken to compensate this smaller engine without increasing the overall cost. That's a complicated question too. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/comment-page-2/#comment-12397 This discussion] confirms the fact that a 1.25 T/W ratio is better costwise. [[File:Gravity_drag_and_thrust_to_weight.png|center]] This graph combines two changing parameters: the overall T/W ratio and the gravity drag that it implies. Changing the gravity drag changes the required delta V, thus changing the mass ratio, that's why the curves have a different minimum. The concerning parameter here is the wet mass, used to design the carrier aircraft. The gravity drag is pretty high compared to what was expected, between 1000 and 1300 m/s in the minimum mass zone (with Isp = 340s). Consequently, we see that the rocket's mass doesn't go below 550 kg. Incidentally, the wet mass doesn't evolve much for T/W ratios between 1.8 and 3.2, and designing a smaller engine because of the cost would not have much bad consequences on the rest of the vehicle. Fixing the overall T/W ratio to 2.5 in our model gives the following rocket: 569.8 kg on the ground, 46.1 kg dry, an engine's thrust of 13.93 kN. Other parameters are [[#Our_first_approximation|above]]. A two-stage rocket should be considered to lower the mass and size of the rocket (3.53 x 0.47 m tanks here) to something more feasible. '''Some questions follow:''' is the engine's T/W ratio constant over the thrust range? How does the structure needs to evolve to cope with the stress of the increasing thrust? How is that correct when considering aerodynamics (drag and lift)? How does the minimum moves when the Isp changes? ==References== <references /> [[Category:Rocket]] 10672fda55cd01805fc0707dad4968f6fc9414ab 479 475 2012-12-29T01:08:25Z Vincent 1 /* category */ changing page name in category wikitext text/x-wiki This page covers the important parameters related to mass for rockets launched from balloon or from aircraft. The goal is to minimize the cost of the vehicle and its carrier aircraft. Minimizing the mass is generally the good way to design a rocket, although it may increase the cost when some parameters are changed, like the rocket's thrust. A trade-off has to be made as explained on this page. Information about particular systems of existing rocket engines can be found in the [[RocketEngines|rocket engines]] page. =First mass approximation for our rocket= The two first sections about [https://en.wikipedia.org/wiki/Delta_v Delta V], combined with the famous [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation], serve as basis to evaluate the [https://en.wikipedia.org/wiki/Mass_ratio mass ratio] of the rocket. Combined with a first evaluation of the mass of [[#Systems_design_for_minimum_mass|each rocket system]], a computer model for rocket mass has been [[#Minimum_mass_evaluation|made]]. The effect of the variation of single parameter on the wet and dry mass of a rocket has been documented as graphs using this model. The gravity drag is a particular parameter affected by other parameters and its variation has dramatic consequences on the mass of the vehicle. It is studied more deeply at the [[#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|bottom of the page]]. ==Delta V requirements== The delta V we will require to achieve orbit depends on several factors: {{SERVER}}/images/formulas_mirror/delta_v_neg.png ===Orbital speed=== The final velocity of the rocket will probably be the same than the required satellite velocity, unless we decide to catapult it somehow. This is the main part of the delta V. At an altitude of 200km, the orbital speed is 7.8 km/s ''for a circular orbit'', and it doesn't change much for an altitude ±100km. We might add some spare speed because if we do only one burn the orbit will not be circular and the delta V will need to be higher. Re-igniting the engine would be required for orbit circularization, but with an SSTO the engine provides so much acceleration on its final seconds that it would require to be very accurate with the burn time. Moreover, the shut-down and ignition procedures introduce losses in propellant mass, which is something we want to avoid. ===Atmospheric drag=== [https://en.wikipedia.org/wiki/Atmospheric_drag#Drag_in_aerodynamics Atmospheric drag] should be very low at this altitude, although it scales badly with small vehicles as explained at many places, in particular in this paper <ref name="francis">Richard J. Francis Jr. '''A systems study of very small launch vehicles'''. In ''Master of Science in Aeronautics and Astronautics at the MIT'', September 1999. ([http://dspace.mit.edu/handle/1721.1/9383 download])</ref>. The paper also models the delta V losses and states that the difference between ground launch and a subsonic 40'000 ft (12km) launch is 1.5km/s, for a 77kg/20cm diameter vehicle, which is huge! (see page 46). ''It still has to be calculated in our conditions.'' If we need to pitch up hardly after rocket release this drag will increase in the beginning of the flight but will quickly reach negligible values. ===Gravity drag=== [https://en.wikipedia.org/wiki/Gravity_drag Gravity drag] is about 1100m/s to 1500m/s for a ground-launch. It is tightly related to the time spent accelerating to the orbital speed, and in fact the only factor if aerodynamic lift is ignored. The longer the vehicle is accelerating, the longer it has to accelerate against gravity, in other words, compensate the gravity acceleration with a 1g vertical acceleration and accelerate vertically to climb above the atmosphere. On the end of the acceleration, when the vehicle reaches an horizontal speed close to the orbital speed, the centrifugal acceleration starts to compensate the gravity. The '''overall thrust-to-weight ratio''' is the vehicle's parameter that affects the time spent burning. It is the thrust-to-weight ratio of the fully loaded vehicle. A ratio of 1 means that the vehicle stands still, if it's vertical. To avoid atmospheric drag losses and minimize gravity drag losses, and from what we could find in the literature <ref name="francis" />, a ratio of 2.0 is common for small vehicles. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/ Another source] states the opposite, that since we're not launching in vertical direction, we don't need as much ratio, and 1.25 is enough. As said above, a larger value implies less loss in gravity drag, so we should opt for a more powerful engine. The issue is then to know which is worse between heavier engine and slower acceleration. This is discussed in the [[#Effects_of_parameter_changes|graphs]] at the bottom. [[File:Rocket_ignition.png|right]] For an aircraft rocket launch, the release velocity will be close to the aircraft's, probably with a pitch of zero (horizontal). The rocket will start falling as soon as it is released. For the rocket engine to compensate this gravity drag, it must be powerful enough and vectored aggressively, for example an overall thrust-to-weight ratio of 1.75 with an thrust pitch of 35 degrees or a ratio of 1.42 with a pitch of 45 deg. Such a high pitch is practically non-realistic as the vehicle is moving forward at the aircraft speed, because it would apply very large stress on the vehicle. Moreover, this just compensates gravity, it's not enough to actually climb. The vehicle would reach Mach 3 before gaining 2km of altitude, which is counter productive in terms of aerodynamic drag (not taken into account in these simulations). This graph depicts the rate of climb for an even more powerful engine and a pitch of 45 degrees. Even if the thrust pitch is 45 deg, the real pitch given by the vehicle's velocity is very slowly going up due to the initial horizontal velocity. The trajectory is very unusual compared to ground- or balloon-launched rockets which start by gaining altitude before horizontal speed. These simple examples show how important it is to use the atmosphere to climb, in the case of an aircraft launched rocket. '''Lift has to be generated to force the vehicle to pitch up quickly and gravity should be partly compensated by lift too''', allowing a less aggressive thrust vector to be used and increasing the gained altitude in the lower rocket flight atmosphere. The atmospheric drag will increase, and this drawback should be evaluated. In the case of Orbital's [https://en.wikipedia.org/wiki/Pegasus_rocket Pegasus] system, it is said [http://colonyfund.com/Reading/papers/phys_econ_leo.html here] that its delta V reduced by approximately 750m/s with the aircraft cruise launch (40'000 feet, 500mph). It uses small wings to help the pitch-up and the climbing, its overall thrust-to-weight ratio is between 2 and 3 and it has 3 stages. From our estimations, ''assuming no aerodynamic drag or lift'', the gravity drag is around 1400m/s when the overall thrust-to-weight ratio is 2.0, and drops below 900 when the ratio is 3.0. ===Rocket release parameters=== If the aircraft can do its release with the same velocity vector than the rocket flight path's beginning, this speed will be completely used by the rocket; if the aircraft needs to release the rocket and then after a few seconds the rocket fires, the z component of the velocity vector is probably lost. Other parameters for the rocket release are: * '''altitude''', the higher it is, the less atmospheric drag and climbing before actually start the useful horizontal delta V. The planned release altitude is 30km ±5km, so that leaves at least 100km to climb, probably even 160km if we want to be able to do 9 orbits, because of atmospheric drag on upper atmosphere for such a small satellite. * [[Flight_at_high_altitude#Approaches_overview|'''high-altitude flight strategy''']], a choice that has to be made between subsonic flight with large wingspan and reduced thrust or supersonic flight with high power engines. ** Choice 1 (subsonic) has a planned horizontal release velocity of 270m/s ±20m/s (0.9M) and a quite small, 5° ± 5° vertical release velocity. ** Choice 2 (supersonic) has a planned release velocity still unknown, possibly 1.5M ±0.3M but has the advantage of giving a higher release altitude and vertical velocity component, likely consistent with the rocket flight path initial vector. The aircraft is even more problematic with this choice. * '''latitude''', important for ground launches, the latitude determines the speed given by Earth's rotation. We have to keep in mind that the aircraft velocity is generally the airspeed, and that the atmosphere is moving to the same average angular speed as the ground. We don't know where we will be able to launch it yet, so let's take a conservative value for now, 300m/s. At equator, it is 465m/s. The fact that's an aircraft or a balloon doesn't really change the value because they don't have an infinite range. ===Delta V evaluation=== Our delta V will then be (if numbers are not explained above, they are just guessed until they are correct): * For subsonic aircraft flight: 7800 (orbital speed) + 80 (atmospheric loss) + 1000 (gravity loss) - 270 (release velocity) - 300 (earth rotation) = 8310m/s * For supersonic aircraft flight: 7800 (orbital speed) + 40 (atmospheric loss) + 700 (gravity loss) - 400 (release velocity) - 300 (earth rotation) = 7840m/s ==Delta V achievement== Once we have a realistic estimation of the required delta V for our rocket, we can start thinking about the characteristics of the rocket by itself and of its engine. The ''simplified'' [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation rocket equation] is expressed as {{SERVER}}/images/formulas_mirror/rocket_equation_neg.png, where ''m₀'' is the initial total mass including propellant (rocket's wet mass), ''m₁'' is the final total mass (rocket's dry mass), and ''v_e'' is the effective exhaust velocity. This simplification is only '''valid for''': no gravity nor aerodynamic drag, constant exhaust velocity (or Isp), ideal nozzle expansion (''p_e'' = ''p₀''), '''and with an initial null velocity''' [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/node103.html]. The calculations and model described below have been done with this simplified equation, yet it is incorrect since in vacuum we have no ideal nozzle expansion and the initial velocity is not null for aircraft-launched rockets. The effect of these two parameters have to be evaluated soon. Below is a list of important parameters for this equation or for the rocket itself, and how they relate to the delta V. * The '''exhaust velocity''' of rocket engine's produced gas is closely related to the '''Isp''' of the engine: {{SERVER}}/images/formulas_mirror/exhaust_velocity_neg.png <ref name="thrust_and_isp">Prof. Z. S. Spakovszky. Unified Thermodynamics and Propulsion, [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/node102.html section 14.1] '''Thrust and Specific Impulse for Rockets'''. MIT teachings, December 2007</ref>. For the duet LOX/E85 we chose, max Isp should be between the max Isp of RP-1 (353s) and max Isp of 92.5% alcohol (338s), so around 344 ± 6s. Isp is [[Rocket:First_approximations#Effects_of_parameter_changes|very important]] and we should aim and actual Isp approaching 10s from the max Isp. Max Isp is vacuum Isp with the most efficient rocket engine and high-expansion nozzle, the sea level Isp is generally around 10% less, and since we will use a cheap design we may not near the optimal efficiency. * '''Thrust''' is related to this exhaust velocity in the following way: {{SERVER}}/images/formulas_mirror/thrust_neg.png <ref name="thrust_and_isp" />. The dotted ''m'' is the mass flow rate of the engine, i.e. how much propellant is burned every second, and the second part can be ignored for approximations. In vacuum, as the external pressure ''p₀'' gets to 0, the larger the nozzle discharge area is, the higher the thrust, but this has a mass [https://en.wikipedia.org/wiki/Rocket_engine_nozzle#Vacuum_use]. Thrust is not directly related to the delta V, but it is related to the mass of the payload that has to be accelerated. For example, [https://en.wikipedia.org/wiki/Ion_thruster ion engines] provide very high Isp and exhaust velocity but their mass flow rate is very low, so is thrust. Therefore, they cannot be used to launch something from Earth to space. * The '''dry mass of the rocket''', once the propellants have been chosen and Isp evaluated, it's the most important factor. The mass ratio (wet mass / dry mass) has a [[Rocket:First_approximations#Effects_of_parameter_changes|strong influence]] on the delta V capability. Conversely, since the mass ratio is fixed by the rocket equation, the lighter the rocket and the more efficient the engine is, the many times less fuel will be required and the smaller the rocket. To make a light rocket, high quality materials, thus expensive, should be used. This is studied in the next section. For the N-prize, the payload mass is nearly insignificant compared to the rocket mass, which is good for the delta V. ** The '''mass of the engine''' itself is closely related to the type fuel delivery system used (what kind of pump or tank pressurization) and to the thrust it produces. The thrust-to-weight ratio depends mostly on chamber pressure, which depends on the capabilities of the fuel pumps system and the efficiency of the regenerative cooling of the chamber and nozzle. Also studied below. As an example, if the delta V is 9.0km/s, and that we have a 340s Isp, the required mass ratio of propellant is 1-e^(-9.0/3.3) = 93.4%, which is very hard to achieve. SpaceX has done it better than 95% [http://spacefellowship.com/Forum/viewtopic.php?f=7&t=11996 (see spacefellowship forum thread on SSTO)] for a first stage, so it doesn't include the payload, fairing, and so on. The [https://en.wikipedia.org/wiki/Haas_(rocket)#Haas_2c Haas 2c] rocket from ARCA (Romania) is also very lightweight and aims to be an SSTO demonstrator. Their [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html engine has] a very nice 110:1 thrust-to-mass ratio, the rocket having a 96.9% mass of fuel. With a 8.0km/s delta V, assuming we have a higher release velocity, the same rocket engine and still only one rocket stage, it becomes 91.1%, still quite hard. A two stage rocket can do this, see the examples in the [https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation#Examples rocket equation Wikipedia page], but if we need to build two rocket engines' systems, we can be sure that we will exceed the money requirement. '''So the main question becomes: on what subsystems can we found cheap alternatives to decrease the mass of the rocket?''' ==Systems design for minimum mass== The mass ratio evaluated above does not tell us what mass we should expect from the rocket. It depends on the minimum mass all systems can be integrated: engine, tanks, structure, avionics, communications and other controls. Again, the paper <ref name="francis" /> addresses this subject. ===Rocket engine=== Since we will have to design our own rocket engine to meet the cost requirement, we can design an engine for any thrust. Sizing the engine is an iterative process: a more powerful or heavier engine will require a much more heavier rocket to achieve orbit because of the mass ratio, but an even more powerful engine is required to propel this new mass and so on. A parameter that we should estimate in the first place is the '''thrust-to-weight ratio''' for the engine, which depends on choices made for the engine design. For example the thickness and material of the engine's wall are linked to the cooling method, material price and manufacturing capability; mass of accessories like pumps and pipes are linked to the pumping method. Once the rest of the rocket's mass has been evaluated, the iterations can be made to have a possible engine mass and final vehicle mass. The thrust-to-weight ratio for the engine doesn't scale linearly either, but in our advantage here <ref name="francis" />. A ratio of more than 100 is common in small launchers' engines <ref name="francis" />, for example 110 for the ARCA [http://spacefellowship.com/news/art29703/arca-has-completed-the-first-executor-rocket-engine.html Executor] engine (ablative cooling, pressure-fed). To minimize cost, we probably will use a conventional pumping system (not turbo-machinery) instead of a tank pressurization system to force the propellants to the engine. This adds weight to the engine (the pump), but removes mass to the tanks and accessories. Our engine's thrust-to-weight ratio will probably be lower than 100. ===Fuel tanks=== ''Main page: [[Rocket_Main_Tank|Rocket tanks]].'' Fuel tanks' mass depends on the type of fuel delivery system used: by tank pressurization or by pumping. When using pressurized tanks to force the fuel into the engine(s), the tanks must have thicker walls to sustain the pressure. When pumping fuel, the tank simply has to prevent it from leaking into vacuum. For kerosene-like fuel, it can be very simple, a plastic tank like in cars. For the cryogenic oxidizer ([[LOX]]), it's more complicated due to metal weakening with large temperature differential and to the evaporation losses related to poor insulation, and may require pressurization anyway. Tanks are the largest part of rockets, they may be used as a structural part too. That requires tanks to be strong enough to overcome the structural stresses on the vehicle. Given the budget, material for the structure and tanks will be aluminium, not titanium and probably not composite materials either. Stainless steel is another possibility but around 3 times more heavy than aluminium. ===Structure=== If tanks are strong enough, they can be used as part of the structure of the rocket. If they are not, and between them in any case, an external skeleton has to be build that will undergo mechanical forces that apply to the vehicle. The ''max Q'', maximum dynamic pressure that the vehicle has to withstand, is reduced when launching from altitude than from ground, so the structure may slightly get lighter. Bending stress is also important as rockets turn or undergo shear winds. Even if we launch from a high altitude, the vehicle should be as aerodynamic as possible. The fairing and the rocket's body must be pretty smooth. If we actually use the tanks as structural components, we may not need an extra outer layer of metal that wraps around the entire rocket, a skirt between them will be enough. The fairing requires a jettison mechanism, the equipment bay requires structure too. ===Avionics and attitude control actuators=== Avionics, including sensors, but excluding actuators and wiring, is a fixed-mass package. Contrary to tanks or structure, it doesn't change when building the more powerful or larger vehicle. <ref name="francis" /> estimated in 1999 that avionics could fit in 50g and 220g for batteries, but it used differential thrust throttling on several engines instead of gimbals or RCS. We will probably have only one engine, and attitude control actuators are mandatory for the rocket's first stage at least. Recent rockets, like ESA's Vega and SpaceX Falcon-1's second stage's use electromechanical, or electromagnetic, actuators. They replace the usual hydraulic actuators (jacks) that require pressurized hydraulic fluid and all their accessories: pumps, tanks, valves and so on. Batteries would have to be upgraded for this use, but off-the-shelf servomotors may be up to this function. Other parts of avionics include the main computing board, the sensors and the wiring/connectors. The computing board can be very small and light-weight but all connectors will take more space and with the wiring will probably be heavier. Sensors mass approximation can be made using readily available components, see the [[EmbeddedRocketComputer#Sensors|sensors page]]. ===Other elements=== Engine's gimbal or RCS, valves, pressurization system (should count as engine mass), communication (should count as avionics), fairing, orbit insertion mechanism, and so on. ==Minimum mass evaluation== A program has been created to make the iterations explained above. The C code, features and version information are available here: [[File:Rocket_mass.c]]. The algorithm is the following: <code> ''start loop ('' compute vehicle '''wet mass''': mass ratio * previous iteration dry mass compute '''propellant mass''' from the difference of wet and dry mass compute '''thrust''' required for the vehicle from the overall thrust-to-weight ratio compute '''fuel and oxidizer mass''' using the stoichiometric ratio and propellant mass, adding more mass for non-storable fuels in case of venting compute '''fuel and oxidizer volumes''' from their mass and density /* now compute the mass of all subsystems */ compute '''tanks''' properties from propellants volumes, pressure, insulation thickness and rocket length-to-diameter ratio compute '''engine mass''' from the engine's thrust-to-weight ratio estimate thrust vector '''actuators''' and battery mass from engine's thrust estimate '''wiring''' mass from the length of the vehicle estimate '''structure''' mass from the size of the vehicle /* prepare next loop */ set next dry mass as the sum of all subsystems listed above PLUS fixed mass systems' mass: '''payload''', '''avionics''' board, '''sensors''', main '''battery''' '') loop n times''</code> An initial dry mass has to be set arbitrarily, the closer it is from the final value, the less iterations are required to converge to it, so it's not really important. The same iteration can be done on wet mass instead of dry mass, the result is the same. ===Our first approximation=== This approximation has been done with the program in version 3. The input values, delta V of 8.3km/s and T/W ratio of 2.5 are the result of [[#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|the analysis]] of the rocket flight with no atmosphere effect on it (drag or lift). We are working on an aerodynamic analysis that may give more realistic values. There are also the values used for the graphs below. '''Input parameters:''' * Mass ratio: 12.0529, from constant Isp=340s and '''DeltaV=8300m/s''' (propellants are 91.7% of the rocket's mass) * Length-to-diameter ratio (tanks): 8 * Engine thrust-to-weight ratio: 100 * Overall thrust-to-weight ratio: '''2.5''' * Payload mass: 0.0199 kg. ''(N-Prize oriented)'' * Avionics (computer and battery) fixed mass: 0.25 kg * Sensors fixed mass: 0.05 kg * Oxidizer: LOX, with 2mm thick aluminium tanks and 10mm thick Aerogel insulation, 1.25kg venting during ascent * Fuel: E85, with 1mm plastic tanks (pumping, no pressurization) and 1mm composite skirt for outer skin * Structure is aluminium (main structure, fairing, aircraft staging mechanism), skirts are composite epoxy/carbon '''Results:''' * DRY MASS: 43.3571 kg, '''WET mass: 522.579 kg''', fairing mass: 3.828 kg * Tanks external size: 3.42936 x 0.461743 m * Engine thrust: 12811.9 N, engine mass: 13.0645 kg * Final acceleration: '''33.0506 G''' * Actuators mass: 2.45561 kg, wiring mass: 0.240055 kg * Structure mass: 3.42766 kg, valves and pipes: 2.26506 kg * LOX tank info: ** diameter: 0.461743 m, length: 2.12526 m ** wall thickness: 0.002 m, volume: 0.294274 m^3, mass: 15.9799 kg * E85 tank info: ** diameter: 0.461743 m, length: 1.3041 m ** wall thickness: 0.001 m, volume: 0.188839 m^3, mass: 1.77612 kg The frost created by cryogenic fluids is not taken into account. The telemetry equipment has not been seriously evaluated yet. ===Effects of parameter changes=== The variation of one parameter is represented in the graphs below, demonstrating the way they affect vehicle's dry and wet mass. Take care about the '''different scales''' of the dry (red) and wet (green) curves. Rocket parameters are the same as above: 12.05 mass ratio (8.3km/s Delta V, 340s constant I_sp), 2.5 T/W ratio, propellants are LOX and E85 at stoichiometric ratio, including 1.25kg of LOX evaporated and tank insulation, full aluminium structure, composite skirts. The software used was version 3 (see updates here: [[File:Rocket_mass.c]]). '''Graphs are valid only for one parameter change''' with the others fixed as above. [[File:Payload_0-80kg.png]] [[File:Payload_0-6k.png]] These two graphs represent the effect of the payload mass to the final mass of the rocket. They use the same data, but the second graph is zoomed-in for very small payloads. There is a slight non-linearity of the vehicle dry and wet mass curves. These graphs also apply to any fixed-mass subsystem, like the avionics or sensors. We can see that for small mass ranges, '''adding 1kg to the vehicle will approximately increase its dry mass of 3.5kg and its wet mass of 45kg''' (remember that's only valid with the input parameters as above). [[File:DeltaV.png]] [[File:Isp.png]] We see above the dramatic effects of delta V increase and I_sp decrease on the rocket mass. A mere 200m/s delta V increase can add 200kg to the wet mass of the rocket, and a too low I_sp makes it simply unrealistic to reach orbit. For an SSTO vehicle, we should try to avoid an I_sp below 340s. The first graph below represents the vehicle mass evolution as function of the mass ratio (wet mass / dry mass). This ratio is given by the I_sp and the required Delta V. The result is consistent with what is explained in the mass ratio [https://en.wikipedia.org/wiki/Mass_ratio Wikipedia's page]. [[File:Mass_ratio.png]] [[File:Engine_thrust_to_weight.png]] The second graph highlights the importance of the engine's thrust-to-weight ratio, fixed by engine's design and the need for high thrust to reduce gravity drag. [[File:Overall_thrust_to_weight.png]] [[File:Final_acceleration.png]] The first graph here depicts the effect of initial acceleration (overall thrust-to-weight ratio) on the vehicle's mass. It related to the gravity drag, the aerodynamic lift, the rocket's trajectory and release velocity, and thus hard to evaluate properly. A higher overall thrust-to-weight ratio gives a more important initial impulse (the acceleration at ignition) and a lower gravity drag, but a heavier engine. In our air-to-orbit scenarios, fortunately, we may benefit from the aircraft release speed and some lift, and the ratio ''may'' be as low as 1.25. To be studied. A possible issue with SSTO vehicles is the final acceleration. Since the engine is sized to propel the fully loaded vehicle and thus a mass much larger than the payload's, when the propellants have been consumed and the fairing jettisoned, the thrust-to-weight ratio is very important. The second graph above shows that even for a small initial acceleration, the final acceleration, if no throttling is implemented, easily reaches 20 or 30 G's. We can see on the graph that the final acceleration varies linearly with initial acceleration, while the engine's thrust required to do so is at best quadratic. ====The gravity drag and overall thrust-to-weight ratio issue==== The [[#Gravity_drag|question]] was: ''is it better to have a more powerful and heavier engine that minimizes the gravity drag because the burn time is decreased or to have a lightweight engine burning longer?'' An issue with this question is the term ''better''. It was first thought as the vehicle with the lowest mass and in this case the answer is in the graph below: it's a compromise. However, when cost is considered, a less powerful engine costs less to manufacture and to design, and it should be evaluated how much more fuel can be taken to compensate this smaller engine without increasing the overall cost. That's a complicated question too. [http://selenianboondocks.com/2008/01/orbital-access-methodologies-part-i-air-launched-ssto/comment-page-2/#comment-12397 This discussion] confirms the fact that a 1.25 T/W ratio is better costwise. [[File:Gravity_drag_and_thrust_to_weight.png|center]] This graph combines two changing parameters: the overall T/W ratio and the gravity drag that it implies. Changing the gravity drag changes the required delta V, thus changing the mass ratio, that's why the curves have a different minimum. The concerning parameter here is the wet mass, used to design the carrier aircraft. The gravity drag is pretty high compared to what was expected, between 1000 and 1300 m/s in the minimum mass zone (with Isp = 340s). Consequently, we see that the rocket's mass doesn't go below 550 kg. Incidentally, the wet mass doesn't evolve much for T/W ratios between 1.8 and 3.2, and designing a smaller engine because of the cost would not have much bad consequences on the rest of the vehicle. Fixing the overall T/W ratio to 2.5 in our model gives the following rocket: 569.8 kg on the ground, 46.1 kg dry, an engine's thrust of 13.93 kN. Other parameters are [[#Our_first_approximation|above]]. A two-stage rocket should be considered to lower the mass and size of the rocket (3.53 x 0.47 m tanks here) to something more feasible. '''Some questions follow:''' is the engine's T/W ratio constant over the thrust range? How does the structure needs to evolve to cope with the stress of the increasing thrust? How is that correct when considering aerodynamics (drag and lift)? How does the minimum moves when the Isp changes? ==References== <references /> [[Category:Rocket|First approximations]] 4aee33525d6b947a50d7ad7320804995beff7c33 6 69 459 445 2012-12-12T01:42:39Z Vincent 1 uploaded a new version of "[[File:Rocket mass.c]]":&#32;clarification about the propellant mass. vehicle_wet_mass is the dry mass + the burnt propellant mass, but it doesn't include the propellant taken aboard and evaporated during the ascent. The output wikitext text/x-wiki ===Description=== This program is an implementation of the iteration required to compute a rocket's mass. It supports single stage rockets (SSTO) with one engine, takes many input parameters such as thrust-to-weight ratios, propellants data, mass ratio or Isp and Delta V, and computes the dry and wet mass of the rocket. Each subsystem's mass is evaluated and the formulas are simple enough to be changed to suit your needs. The rocket is modelled as depicted below: [[File:Rocket_model.png|700px]] The fairing jettison is currently (v3) not accounted for in the delta V calculation. It is made of aluminium, so it's generally quite heavy, even if it's thin. The pipes and valves have been added in v3 too, also adding an important mass that was forgotten in previous versions. This result in the rocket being around 150kg heavier, because each kg of rocket has approximately 50kg of propellant needed to lift it. The structure's mass has been improved too, taking into account solid tanks for the structure. ===Features=== See the beginning of the file for all input parameters. Each subsystem has a function computing its mass given the rocket length or other parameters, they can be easily modified. See [[Rocket:First_approximations#Our_first_approximation|first approximation]] for the list of output parameters and an example. The mass ratio can be either given or computed from I_sp and Delta V. The formula used is very simple but assumes the exhaust velocity (or I_sp) to be constant [https://en.wikipedia.org/wiki/Rocket_equation#Examples (source)]: e^(-DeltaV/V_e), V_e being the gas exhaust velocity, which can be calculated from Isp: V_e = Isp * g₀. [[File:Rocket_mass_tanks_v1-2.png|right|300px]] The dimensions of the tanks is a major issue. In version 1 and 2, both tanks are computed with the same inner diameter. The goal is to have them roughly the same size so that the rocket can be a simple cylinder. In version 2, tank insulation has been added for the mass concern only. [[File:Rocket_mass_tanks_v3.png|right|300px]] In future version (coming soon) the insulation and wall thickness will be taken into account in the ''outer'' diameter of tanks, so that they can really fit in the same cylinder in any case. Insulation is Aerogel, 13kg/m³ density. The insulation thickness can be calculated from heat and thermal transmission data using this other program: [[File:Heat_transfer_to_rocket_tank.c]]. Tank wall thickness is not properly computed from the pressure and heat, but only estimated as 1mm per 10 atmospheres of pressure, with a minimum value of 1mm for pressurized tanks. The storable fuel tank is currently made of plastic to reduce weight, since pumps will probably be used instead of pressurized tanks. Structural parts are made of aluminium, except the ullage (v2) and soft tank skirts (v3) which is a carbon/epoxy composite. The vehicle's frame is composed of 5 profiled rods, 2mm thick, 15mm wide. This ''may'' be enough to support the aircraft staging mechanism and the rocket's weight and flight stress. ===Compilation=== Compile with <code>gcc -O2 -Wall rocket_mass.c -o rocket_mass -lm</code> for normal use, and add <code>-DMAKE_GRAPH</code> after gcc on the command line if you want to create graph data as visible on the [[Rocket:First_approximations#Effects_of_parameter_changes|first approximations page]]. 1746c486ff0a4aa003f76fad82a9f8262759aa76 6 81 460 2012-12-12T03:05:29Z Vincent 1 Made with [[File:Rocket_mass.c]] v3 plus a gravity drag evaluation added in the iteration. wikitext text/x-wiki Made with [[File:Rocket_mass.c]] v3 plus a gravity drag evaluation added in the iteration. fc5f8056505940ac8a2a7658d513febdfcc6d5f7 461 460 2012-12-12T03:33:23Z Vincent 1 uploaded a new version of "[[File:Gravity drag and thrust to weight.png]]":&#32;Updated with the gravity drag curve and make the image larger. wikitext text/x-wiki Made with [[File:Rocket_mass.c]] v3 plus a gravity drag evaluation added in the iteration. fc5f8056505940ac8a2a7658d513febdfcc6d5f7 0 1 463 440 2012-12-12T15:46:47Z Vincent 1 /* News */ about the gravity drag issue wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''December 2012 update:''''' A simple rocket trajectory model has been made in order to evaluate the trajectory of an aircraft- or balloon-released-rocket. It appears that the Delta V taken for granted for gravity drag for these rockets, around 800 m/s, is quite erroneous, or not possible with the expected overall thrust-to-weight ratio. Without taking into account the aerodynamic effects like lift and drag, for a ratio of around 1.5, the Delta V for gravity drag is at least 1300 m/s. To reduce it, the ratio should be higher, like 3.0, in that case it may be possible to have only 800 m/s Delta V for gravity, but the mass of the engine would be much higher. See [[Rocket:First_approximations#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|a solution]] to this issue. The program evaluating the mass of the rocket has been refined (v3), and the forgotten pipe linking the upper tank to the engine adds approximately 150 kg of wet mass too ([[File:Rocket_mass.c]]). Incoming works will focus on how much we can expect the aerodynamic lift to compensate the gravity drag, and how much we can reduce the engine's thrust to decrease the mass of the vehicle while staying air-SSTO. '''''November 2012 update:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 150 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] have been created to illustrate the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. The program ([[File:Rocket_mass.c]]) modelling the mass of rockets is being refined and the cryogenic propellant vaporization has been evaluated by another program ([[File:Heat_transfer_to_rocket_tank.c]]). Both programs are freely available. '''''August 2012 update:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. '''''November 2011 update:''''' Information available on this site is sometimes outdated, and may be weakly verified or partly false information, since it was done with little knowledge on the topics at the time. A documentation base is being built to provide access to all or a major part of information used to develop the project; the website pages are slowly updated to reflect the actual progress. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. See [http://gravityloss.wordpress.com/2008/04/21/air-breathers-advantage this article] for more information on the differences of rocket and aircraft propulsion efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. [http://www.youtube.com/watch?v=esgc5W_Ufng This one] (SpaceX guys), here captured at SpaceUP, doesn't even allow attitude control out of atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Xcor has created in 2003 a [http://filespump.info/piston_pumps.html piston pump] for LOX, which is now used on a 1,500 lb-thrust LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory interpolation is closely tied to [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control]. '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. 6079f347ee8ecd047967b25c97a4736495ff297b 485 463 2012-12-29T01:43:20Z Vincent 1 text updates, more links, more fixes wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''December 2012 update:''''' A simple rocket trajectory model has been made in order to evaluate the trajectory of an aircraft- or balloon-released-rocket. It appears that the Delta V taken for granted for gravity drag for these rockets, around 800 m/s, is quite erroneous, or not possible with the expected overall thrust-to-weight ratio. Without taking into account the aerodynamic effects like lift and drag, for a ratio of around 1.5, the Delta V for gravity drag is at least 1300 m/s. To reduce it, the ratio should be higher, like 3.0, in that case it may be possible to have only 800 m/s Delta V for gravity, but the mass of the engine would be much higher. See [[Rocket:First_approximations#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|a solution]] to this issue. The program evaluating the mass of the rocket has been refined (v3), and the forgotten pipe linking the upper tank to the engine adds approximately 150 kg of wet mass too ([[File:Rocket_mass.c]]). Incoming works will study aerodynamic lift and drag of a simple rocket and of a rocket with small supersonic wings to better evaluate the gravity drag for our mass evaluation. '''''November 2012 update:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 150 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] have been created to illustrate the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. The program ([[File:Rocket_mass.c]]) modelling the mass of rockets is being refined and the cryogenic propellant vaporization has been evaluated by another program ([[File:Heat_transfer_to_rocket_tank.c]]). Both programs are freely available. '''''August 2012 update:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. See [http://gravityloss.wordpress.com/2008/04/21/air-breathers-advantage this article] for more information on the differences of rocket and aircraft propulsion efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several N-Prize teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. In [http://www.youtube.com/watch?v=esgc5W_Ufng this video] (SpaceX guys), here captured at SpaceUP, they don't even predict the use of attitude control outside the atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell [http://www.jpaerospace.com/] is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) nearly equivalent to the empty weight of the plane (20,000kg) and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]], Category page: [[:Category:Rocket|list of pages]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost, but also regarding their efficiency, the Isp they can produce. This is [[Rocket:First_approximations#Effects_of_parameter_changes|especially important]] in an SSTO design. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Turbopumps are used on full-scale commercial rockets, but are very complex to build and design. Xcor has created and demonstrated since 2003 a [http://www.xcor.com/products/pumps/ piston pump] for LOX, which is now used on a 1,500 lb-thrust (6.6 kN) LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory is closely tied to the [[Flight_at_high_altitude#Approaches_overview|initial release parameters]], the [[Rocket:First_approximations|flight parameters]] such as thrust and aerodynamics, and [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control], which depends on sensors and actuators: '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. 4bbcfa918eaf8606ff4758c99509d9698fc7b4f2 494 485 2013-01-11T00:46:29Z Vincent 1 /* The aircraft */ Bladon link wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''December 2012 update:''''' A simple rocket trajectory model has been made in order to evaluate the trajectory of an aircraft- or balloon-released-rocket. It appears that the Delta V taken for granted for gravity drag for these rockets, around 800 m/s, is quite erroneous, or not possible with the expected overall thrust-to-weight ratio. Without taking into account the aerodynamic effects like lift and drag, for a ratio of around 1.5, the Delta V for gravity drag is at least 1300 m/s. To reduce it, the ratio should be higher, like 3.0, in that case it may be possible to have only 800 m/s Delta V for gravity, but the mass of the engine would be much higher. See [[Rocket:First_approximations#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|a solution]] to this issue. The program evaluating the mass of the rocket has been refined (v3), and the forgotten pipe linking the upper tank to the engine adds approximately 150 kg of wet mass too ([[File:Rocket_mass.c]]). Incoming works will study aerodynamic lift and drag of a simple rocket and of a rocket with small supersonic wings to better evaluate the gravity drag for our mass evaluation. '''''November 2012 update:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 150 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] have been created to illustrate the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. The program ([[File:Rocket_mass.c]]) modelling the mass of rockets is being refined and the cryogenic propellant vaporization has been evaluated by another program ([[File:Heat_transfer_to_rocket_tank.c]]). Both programs are freely available. '''''August 2012 update:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012 update:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012 update:''''' Study of aerodynamics is under way. More man power is expected in April. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. See [http://gravityloss.wordpress.com/2008/04/21/air-breathers-advantage this article] for more information on the differences of rocket and aircraft propulsion efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several N-Prize teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. In [http://www.youtube.com/watch?v=esgc5W_Ufng this video] (SpaceX guys), here captured at SpaceUP, they don't even predict the use of attitude control outside the atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell ([http://www.jpaerospace.com/ JP Aeroospace]) is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) equivalent to the empty weight of the plane and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Small gas turbine engines exist however, even in an axial design, see [http://www.bladonjets.com/technology/gas-turbines/ Bladon Jets]. Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]], Category page: [[:Category:Rocket|list of pages]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost, but also regarding their efficiency, the Isp they can produce. This is [[Rocket:First_approximations#Effects_of_parameter_changes|especially important]] in an SSTO design. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Turbopumps are used on full-scale commercial rockets, but are very complex to build and design. Xcor has created and demonstrated since 2003 a [http://www.xcor.com/products/pumps/ piston pump] for LOX, which is now used on a 1,500 lb-thrust (6.6 kN) LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory is closely tied to the [[Flight_at_high_altitude#Approaches_overview|initial release parameters]], the [[Rocket:First_approximations|flight parameters]] such as thrust and aerodynamics, and [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control], which depends on sensors and actuators: '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. ae9dfe19b409e5ed80345fd3724f94be1345b7f7 8 2 468 107 2012-12-13T02:01:15Z Vincent 1 trying to put subsections wikitext text/x-wiki * navigation ** mainpage|mainpage ** Resources|Resources ** recentchanges-url|recentchanges ** randompage-url|randompage * sections ** EmbeddedRocketComputer|Computer ** Build_a_cheap_turbofan|Plane engine ** RocketEngines|Rocket engine *** Rocket:First_approximations|Mass approximation *** Rocket_Main_Tank|Tanks ** Testing|Testing and validation f41460ec9f5fbec1da063d36a8b7792ac9cbe015 469 468 2012-12-13T02:01:53Z Vincent 1 reverting: dispayed as sections wikitext text/x-wiki * navigation ** mainpage|mainpage ** Resources|Resources ** recentchanges-url|recentchanges ** randompage-url|randompage * sections ** EmbeddedRocketComputer|Computer ** Build_a_cheap_turbofan|Plane engine ** RocketEngines|Rocket engine ** Testing|Testing and validation 3cde20a1dd597668d5fc7f08ecb808361e649a2d 14 82 476 2012-12-29T00:58:29Z Vincent 1 rocket category wikitext text/x-wiki The Rocket category groups pages related to rocket systems design, flight environment and parameters, design and modelling. 9807260c69830ec63d95cf8150edcd7f3f3b45d6 0 17 477 437 2012-12-29T01:01:59Z Vincent 1 Adding rocket category links wikitext text/x-wiki This page is an introduction to rocket engines, explaining and comparing some of its subsystems. Our rocket will be based on regular rocket technology, as explained on this page. However some choices need to be made in order to gauge the feasibility of the project as a whole, in particular to have ideas of the possible dimensions of the rocket. These choices are presented on the page of the [[Rocket:First approximations|first approximations for the rocket]]. Other information and pages about the rocket and its flight can be found in the [[:Category:Rocket|Rocket category]]. =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, instead of fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine failure. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotary Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Using LOX for film cooling has also been [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=369#EngineDevelopment demonstrated], by Armadillo Aerospace. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. From SpaceX's experience, it seems that regenerative cooling is sufficient to sustain the combustion temperature hitting the walls. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. [[Category:Rocket]] 54d1b7cd39a5814df5271d65f6b01e3688b84e82 478 477 2012-12-29T01:08:18Z Vincent 1 /* category */ changing page name in category wikitext text/x-wiki This page is an introduction to rocket engines, explaining and comparing some of its subsystems. Our rocket will be based on regular rocket technology, as explained on this page. However some choices need to be made in order to gauge the feasibility of the project as a whole, in particular to have ideas of the possible dimensions of the rocket. These choices are presented on the page of the [[Rocket:First approximations|first approximations for the rocket]]. Other information and pages about the rocket and its flight can be found in the [[:Category:Rocket|Rocket category]]. =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, instead of fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine failure. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotary Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Using LOX for film cooling has also been [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=369#EngineDevelopment demonstrated], by Armadillo Aerospace. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. From SpaceX's experience, it seems that regenerative cooling is sufficient to sustain the combustion temperature hitting the walls. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. [[Category:Rocket|General presentation]] 967fc4ba2dea6d7846537975b783aba5c33e68e8 480 478 2012-12-29T01:09:10Z Vincent 1 /* category */ changing page name in category wikitext text/x-wiki This page is an introduction to rocket engines, explaining and comparing some of its subsystems. Our rocket will be based on regular rocket technology, as explained on this page. However some choices need to be made in order to gauge the feasibility of the project as a whole, in particular to have ideas of the possible dimensions of the rocket. These choices are presented on the page of the [[Rocket:First approximations|first approximations for the rocket]]. Other information and pages about the rocket and its flight can be found in the [[:Category:Rocket|Rocket category]]. =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, instead of fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine failure. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotary Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Using LOX for film cooling has also been [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=369#EngineDevelopment demonstrated], by Armadillo Aerospace. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. From SpaceX's experience, it seems that regenerative cooling is sufficient to sustain the combustion temperature hitting the walls. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. [[Category:Rocket|Rocket engines in general]] 8b3120a5ada67d214b797f42fdbcf5f8a3c6e0d2 0 75 481 428 2012-12-29T01:12:13Z Vincent 1 adding rocket category wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = Nu The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ===The Rayleigh number ''Ra''=== The [https://en.wikipedia.org/wiki/Rayleigh_Number Rayleigh number] is an indicator for natural convection heat transfer. When it is below the critical value for a fluid, heat transfer is primarily in the form of conduction; when it exceeds the critical value, heat transfer is primarily in the form of convection. {{SERVER}}/images/formulas_mirror/rayleigh_number_neg.png α is the [https://en.wikipedia.org/wiki/Thermal_diffusivity thermal diffusivity] {{SERVER}}/images/formulas_mirror/thermal_diffusivity_neg.png ; β is the [https://en.wikipedia.org/wiki/Coefficient_of_thermal_expansion thermal expansion coefficient], for an isobaric process it can be approximated to β = 1/T. ''x'' is the distance from the leading edge. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of cryogenic propellant on the ground and during the ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. The program using the [[Flight_at_high_altitude#Gas_properties_and_altitude|atmospheric model]] and the flight path model to compute the heat transfer during flight is available freely here: [[File:Heat_transfer_to_rocket_tank.c]]. '''''Warning:''''' ''the values presented below are highly dependent on the dimensions of the considered tank (diameter: 0.27m, length: 1.63m, surface: 1.38261m^2), which was given by the [[Rocket:First_approximations#Minimum_mass_evaluation|first approximation]] of the rocket mass program, the climb rates and flight profile (currently based on [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2]'s rate), and the temperature of the propellant (here LOX at 92K).'' '''''Warning:''''' ''the method and results presented here are based on approximations and assumptions, and it may even have not been done in the proper way. Please validate the method used if you have some knowledge on convective heat transfer.'' ===Natural convection for horizontal cryogenic tank=== On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. {{SERVER}}/images/formulas_mirror/h_natural_conv_external_horiz_cyl_neg.png D is the diameter of the approximated cylinder, in our case we take 0.27m, given by a [[Rocket:First_approximations#Minimum_mass_evaluation|first evaluation]] of the rocket mass program. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C or 298.15 K ground temperature: * the density (rho) is 1.18391 kg/m^3 and at the mean film temperature of -78.07°C or 195.075 K, given the propellant and tank temperature of 92 K: * the viscosity µ is 1.30015e-05 Pa.s (calculated using [http://www.cfd-online.com/Wiki/Sutherland's_law Sutherland's law]) * the kinematic viscosity nu is µ / rho = 1.0982e-5 m^2/s * the specific heat Cp is 1007.68 J/kg.K (approximated using [http://ninova.itu.edu.tr/tr/dersler/ucak-uzay-fakultesi/965/uck-421/ekkaynaklar?g96162 this interpolation]) * the Prandtl number is 0.73862 * the thermal conductivity of air ''k'' is 0.0177375 W/m.K (approximated using [http://physics.tutorvista.com/heat/heat-transfer.html#thermal-conductivity-of-air this interpolation]) * α is thus k/(rho.Cp) = 1.4868e-5 m^2/s * β is approximated to 1/T = 1/195.075 = 5.126e-3 1/K * Ra_D is ((9.80665 * 5.126e-3) / (1.0982e-5 * 1.4868e-5)) * (298.15 - 92) * 0.27^3 = 6.3467e10 * 0.27^3 = 1.24931e+9. * finally, ''h'' is 0.0177375/0.27 * (0.6 + (0.387 * 1.24931e+9^1/6) / (1 + (0.559/0.73862)^9/16)^8/27)^2 '''= 8.20533 W/m^2.K''' We finally have a heat transfer coefficient for air. We can compute the total heat transferred from it, using Newton's law equation [[Heat_transfer#Heat_transfer_coefficient_h|(top)]]. We need to specify the area on which the heat transfer will apply, the approximated cylinder, and the number of seconds during which the heat is transferred. Tanks are a cylinder with two hemispherical end-caps. Since the end-caps still conduct heat, we will include their area and assume it is part of the approximed cylinder's area. Assumed area is then 1.3826m^2. Let's take 600 seconds (10 minutes) for the time spent on the ground between tanks filling and aircraft lift-off. We assume the temperature to be constant on the ground and in the tank while heat is transferred. This is correct since the heat energy feeds a phase change in the propellant (vaporization) and not a temperature increase of some sort. ''dQ/dt'' = ''h'' * ''A'' * (''T_ground'' - ''T_prop'') = 8.20533 * 1.3826 * (298.15 - 92) = 2338.73 J/s. We can now [[Rocket_Main_Tank#Calculating_evaporation_rate|calculate]] the evaporation rate of the propellant with and without insulation. The heat of vaporization for [[LOX]] is 213 kJ/kg. Without insulation, the vaporization rate of LOX is 11g/s, making 6.59kg of LOX evaporated in 600s. With a 10mm [[Rocket_Main_Tank#Thermal_insulation_materials|Aerogel]] insulation, the thermal conductivity being 15 mW/m.K, its U-value being k/L = 0.015/0.010 = 1.5 W/m^2.K, the vaporization rate is 919.5mg/s. For 10 minutes, it goes down to 0.552kg (instead of 6.59 without insulation). The overall heat transfer coefficient U was calculated using [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#Alternative_Method_.28A_simple_method_for_determining_the_overall_heat_transfer_coefficient.29 this] method. ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2], ''Heat transfer in turbulent flow over a flat plate'', is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr varies between 0.73 and 0.75, and the Reynolds number for the characteristic length L of the tank (Re_L) drops from 4.36714e+07 to 6.68455e+06 with altitude. This value of Re confirms that the boundary layer is turbulent. [2] indicates that the Colburn analogy can still be used for Pr up to 10^8, but with some loss of accuracy. We use the following approximation to calculate the average Nusselt number for forced convection in turbulent flows: Nu_L = 0.0366 (Re_L^0.8 - 14500) Pr^1/3 given in [2]. The calculated average Nusselt number for ascent operation is 42353 at sea level and drops to 1037 at 30km altitude. ''h_L'' can then be calculated too, ''h_L'' = Nu_L * ''k''_air / L. It varies from 460.8 on the ground to 9.6 W/m^2.K at 30km altitude. The simple flight path approximation is used to calculate the average ''h'' value for every second of the flight. We also need to calculate the dQ/dt value for each second, since the external temperature changes as the aircraft ascends. The integrated value of heat Q is finally: Q = 1710.94 kJ. '''For [[LOX]], it translates into more than 80kg of oxidizer evaporated during the ascent, when no insulation is used. With an Aerogel insulation layer of 10mm, the mass of evaporated LOX is evaluated to 0.666kg.''' That corresponds to Q = 141.89 kJ. The overall heat transfer coefficient U was calculated using [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#Alternative_Method_.28A_simple_method_for_determining_the_overall_heat_transfer_coefficient.29 this] method. ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 3. Lecture on '''Advanced Heat Transfer''', Ibrahim Sezai, 2011. Distributed as open courseware, [http://opencourses.emu.edu.tr/course/view.php?id=23 here] [[Category:Rocket|Heat transfer]] 6f19cab06871334182755d57e2a6aa72497088e7 0 71 482 379 2012-12-29T01:12:38Z Vincent 1 Adding rocket category wikitext text/x-wiki =Propellants= Here are some links: [http://www.thespacerace.com/forum/index.php?topic=2583.0 tables and info]. Wikipedia's [https://en.wikipedia.org/wiki/Liquid_rocket_propellants#Bipropellants table]. Another [http://www.braeunig.us/space/propel.htm table]. ==Fuels== * RP-1, storable, not dangerous, available?, moderately cheap, density: 806 kg/m³ * E85, storable, not dangerous, readily available, cheap, density: 780 kg/m³ * LH₂, cryogenic, not dangerous, requires specific storage and permit?, quite expensive, density: 71 kg/m³ * Ethanol, storable, not dangerous, readily available, cheap, density at 92.5%: 800 kg/m³ ==Oxidizers== Besides the cryogenic issue, LOX is probably the safest oxidizer. Others may be storable at nearly ambient temperature, or under pressure, but they are less stable, subject to explosion or toxic, and more expensive. In a cold country, Nitrous oxide can be a good alternative, but if temperature is around 20°C its density is too low and tanks require a large spare volume. Nitrous oxide is at least 20 times more expensive than LOX too. * [https://en.wikipedia.org/wiki/Liquid_oxygen LOX] ([[LOX|internal link]]), cryogenic, explosive, requires cryogenic storage, cheap, density: 1141 kg/m³ at 92.2K and 1 atm, 974.42 kg/m³ at 120K and 10bar * [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] (N₂O), refrigerated liquid (boiling at -88.5°C) or self pressurizing (vapour pressure at 20°C is ~50.1 bar), but critical point is 36.4°C and 72.45 bar, non-toxic, quite expensive, density: 1223 kg/m³ at -88.5°C, 750 kg/m³ at 20°C, changes dramatically with temperature * [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide], pressurized, self-decomposes explosively, expensive to have it manufactured at a high concentration, density: 1450 kg/m³ (pure) * [https://en.wikipedia.org/wiki/Nitrogen_tetroxide Nitrogen tetroxide] (N₂O₄), storable, highly toxic, density: 1443 kg/m³ at 21°C. ==Mixes== The stoichiometric ratio is not the optimal ratio for rocket applications. The density and exhaust speeds are more important than maximum temperature. In particular, the ratio T_c/M, combustion temperature / molecular mass, is a good indicator of the exhaust speed, as explained [http://www.thespacerace.com/forum/index.php?topic=2583.msg17485#msg17485 here]. {| border="1" class="wikitable" !Oxidizer !Fuel !I_sp (sea lvl) !max I_sp (vacuum) !Stoichiometric !T_c Combustion temp. (K) !Average density (kg/l) |- |rowspan="4"|LOX |LH₂ |381 (r=5.0) |455 |8 |3304 |0.32 |- |RP-1 |289 (r=2.29) |353 |2.56 |3526 |1.02 |- |E85 |? |? |2.26 |around 3360 |around 1 |- |Ethanol 95% |277 |? |2.19 |3314 |0.97 |} [[Category:Rocket|Propellants]] be9ba02de91e48cca1f8e34e75e4fc0e2dceb6a8 0 62 483 383 2012-12-29T01:13:06Z Vincent 1 Adding rocket category wikitext text/x-wiki =Liquid Oxygen= ''See [[Propellants]] page and [https://en.wikipedia.org/wiki/Oxygen Wikipedia's page]. See also [[Rocket_Main_Tank#Cryogenic_fuel_tanks|tank page]] for insulation.'' Boiling point of -182.95°C (90.20 K). Critical temperature of -118.59°C (154.59 K) meaning that it cannot stay liquid above this temperature. Critical pressure is 5.043 MPa (49.77 atm). Density: 1141 kg/m3 at 92.2K and 1 atm, 974.42 kg/m3 at 120K and 10bar 1 Litre of LOX provides 840 Litre of GOX (gaseous oxygen) at 1atm/20°C. Like any liquid, LOX vaporizes when stored above its boiling point. The vapour builds up to pressure that the tank will no be able to contain, so a relief valve must be put in place to compensate this evaporation. Insulation limits the rate of vaporization, the best being the vacuum space between a tank inner and outer walls, like Dewar flasks. Latent heat of vaporization: 213 kJ/kg (6.82 kJ/mol). [[Category:Rocket|LOX]] d679ed81c1968854e3ea7778428f5fe28153a3b3 490 483 2013-01-05T22:49:36Z Vincent 1 /* Liquid Oxygen */ safety link wikitext text/x-wiki =Liquid Oxygen= ''See [[Propellants]] page and [https://en.wikipedia.org/wiki/Oxygen Wikipedia's page]. See also [[Rocket_Main_Tank#Cryogenic_fuel_tanks|tank page]] for insulation.'' Boiling point of -182.95°C (90.20 K). Critical temperature of -118.59°C (154.59 K) meaning that it cannot stay liquid above this temperature. Critical pressure is 5.043 MPa (49.77 atm). Density: 1141 kg/m3 at 92.2K and 1 atm, 974.42 kg/m3 at 120K and 10bar 1 Litre of LOX provides 840 Litre of GOX (gaseous oxygen) at 1atm/20°C. Like any liquid, LOX vaporizes when stored above its boiling point. The vapour builds up to pressure that the tank will no be able to contain, so a relief valve must be put in place to compensate this evaporation. Insulation limits the rate of vaporization, the best being the vacuum space between a tank inner and outer walls, like Dewar flasks. Latent heat of vaporization: 213 kJ/kg (6.82 kJ/mol). [http://www.nasa.gov/centers/wstf/laboratories/oxygen/index.html About] safety and LOX systems, NASA White Sands Test Facility. [[Category:Rocket|LOX]] acefbd0e3e1a519608e16f26e75c3ff6034666c2 0 24 484 425 2012-12-29T01:13:43Z Vincent 1 adding rocket category wikitext text/x-wiki =Rocket Fuel tanks= [[Image:Restone_tank.jpg|right|320px]] In modern launchers, two tanks are used, one for fuel and one for the oxidizer, but it has not always been the case. The Redstone rocket for example used a single tank with an internal separation, as we can see below. The sphere is the most lightweight volume (volume / area is minimized), but we can't have rockets as large as they are long, so cylinders with hemispheric caps are used. Having a single tank cut in two like for the Restone is efficient mass-wise and volume-wise but can bring new issues for insulation in case of a single cryogenic fluid (funny enough, that was the case for the Restone which used LOX and ethanol). The thicker insulation may overtake the mass benefits of a single tank. ==Sloshing and other effects== Special care must be taken to avoid sloshing and vortexes in the tanks, that may lead to bubbles in propellant flow. ==Wall thickness and material== Tank material first has to be stable with what's inside. Lists are available for cryogenic liquids at least. Besides this basic filter, the material choice mostly depends on money and on what's available on the market. For pressurized tanks, we will use 6061 aluminium or steel. The thickness of the tank walls obviously depend on the internal pressure, but also on the diameter of the tank. See [http://www.innovatia.com/Design_Center/FundRoc_4-8.htm]. For example, aluminium walls can be 2mm thick and 0.4m wide for a pressure up to 13 bar with no safety factor. For a 0.2m wide tank, the thickness can be 1mm for the same pressure, or twice the pressure for the same thickness. ==Cryogenic fuel tanks== Cryogenic and also low boiling temperature liquids like nitrous oxide are persistently evaporating at ambient temperature. It's like having water at 100°C and providing always more heat to it. When the vapour pressure is high enough, and when tanks are solid enough too, the evaporation can reach [https://en.wikipedia.org/wiki/Evaporation#Evaporative_equilibrium equilibrium] and the tank can contain a stable mix of liquid and gas at high pressure. This is the case for nitrous oxide at temperatures below 36.4°C, its critical temperature above which it turns all into gas, no matter what pressure is used. The issue then becomes the density of the mixture, which drops greatly. For [[LOX]], the critical temperature is -118.59°C, and the critical pressure is 50.43 bar. There's no point in keeping it so much pressurized because it could just boil off at this temperature. Since the phase change occurs at a constant temperature, we can as well choose a temperature and a pressure at which the LOX density is high enough, but that's a trade-off with the evaporation rate. Since the temperature difference between inside and outside the tank is greater, even more heat is transferred to the LOX, and more evaporation is created. Tank insulation is then required to avoid venting all the propellant before actually using it (balloon or aircraft launch can take some time to get to the launch altitude). ===Calculating evaporation rate=== ''Main page: [[Heat transfer]]'' In the case of balloon or aircraft launches, the rocket stays filled with cryogenic fluids for a quite long time before being fired, with no possibility for refuelling. First on the ground after tank filling, during the final pre-flight verifications, then during the ascent to the rocket ignition altitude. The vehicle passes through different air layers with different temperatures and densities. A simple [[Flight_at_high_altitude#Gas_properties_and_altitude|atmospheric model]] has been made (see [[File:ISA_atmospheric_model.c]]) and is used in order to calculate the heat transferred to the vehicle during the ascent, which is dominant over the heat transferred during the powered flight of the rocket. The ascent rate is approximated, as constant between some gross altitude levels. The evaporation rate can be calculated easily once the amount of heat provided by the surrounding air is known, using the heat of vaporization ΔH_vap of the propellant, the insulation layer's U-value, temperatures and so on. The issue is that the amount of heat provided by air is very hard to calculate. It should be done using CFD, but can be approximated in some conditions. The convection heat transfer coefficient ''h'' is the key concern. See the [[heat transfer]] page for more information and '''results''' for the LOX tank given by the first rocket mass [[Rocket:First_approximations#Our_first_approximation|approximation]]. ===Thermal insulation materials=== Cryogenic fuel tanks benefit from being insulated, which limits vaporization or even prevents boiling. A list of thermal conductivities is available on [https://en.wikipedia.org/wiki/List_of_thermal_conductivities Wikipedia]. PU foam is a simple solution, it can be sprayed at the desired thickness but may be hard to spray in very thin layers. Aerogel is the best existing insulation material and it has been used by Armadillo Aerospace (see [[Rocket_Main_Tank#Resources|below]]). Aerogel for cryogenic applications is [http://www.aerogel.com/products/overview-product.html currently available] in 5mm or 10mm thick sheets. {| border="1" class="wikitable" !Material !''k'' (mW/m.K) !density (kg/m³) !availability,comments |- |Aerogel |15 |13 |readily available in sheets, cheap |- |PU foam |22 |a density of 24 to 32 (1.5 to 2 LB/cu.ft) should be enough |readily available, cheap, sprayed |- |Expanded polystyrene |32 to 38 |40 to 15 (resp.) |readily available in boards, cheap |- |cotton |around 30 | |readily available, cheap |- |mineral insulation |around 40 | |readily available, cheap |- |neoprene |54 |960 |readily available, cheap, heavy |} A more precise list of low conductivity materials is available [https://en.wikipedia.org/wiki/List_of_insulation_material here]. ==Propellant lines== Pumps and Engine fuel supply pipe and valve, tank pressure sensor, fill and drain pipes and valves. For a cryogenic fuel or a high vapour pressure fuel tank: pressure relief valve, venting valve. ==Resources== Armadillo Aerospace has [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=333 documented] their LOX tank insulation: We settled on using Pyrogel insulating blankets from Aspen Aerogels to insulate our tanks: [http://www.aerogel.com/products/pdf/Pyrogel_6250_DS.pdf#search=%22pyrogel%20insulation%22] After Phil figured out the right pattern to make the gores for the spherical tanks, it went very well. We used a spray adhesive to attach it, and we tested all the combinations of dusty side / non-dusty side and painted / non-painted for best adhesion. Surprisingly, putting the adhesive on the dusty side worked best. The material still does shed some dust in the wind, but it is reasonably rugged, more so than the fastblock insulation we were previously using, and it is only about $5 / square foot, which is a tenth the cost of the fastblock. It also sheds water fairly well. I was a bit surprised at how much of a difference insulating the tank made on our boiloff rates. I had been presuming that much of the boiloff was due to heat conduction from the rest of the 90 pound tanks that don’t get cooled that well during filling, but it turns out that the entire tanks get cooled a lot better with insulation. We used to have 10 psi in the tanks after filling, even with the vents open, but now it is just 3 psi and soon drops to 1 psi. We also insulated our test stand tank. http://media.armadilloaerospace.com/2006_08_12/insulating.jpg XCOR uses Styrofoam, a type of polystyrene with a 33 mW/m.K thermal conductivity, to insulate their aluminium LOX tanks. [[Category:Rocket|Tank]] fed67778343adde476d2bcf7923a1cb10907d42c 0 83 486 2013-01-01T02:32:24Z Vincent 1 Page creation, context and first document link wikitext text/x-wiki =Rocket aerodynamics= We have seen with the [[Rocket:First_approximations|first approximation]] that the gravity drag is higher than expected for a single stage air-to-orbit configuration, when aerodynamic effects are ignored. The rocket engine has to be larger to compensate the gravity. On this page, we will evaluate how the aerodynamic effects can be used to compensate the gravity ([https://en.wikipedia.org/wiki/Lift_(force) lift]) without impacting the thrust too much ([https://en.wikipedia.org/wiki/Aerodynamic_drag drag]). This will be evaluated for a rocket without wings or fins at first, then we will do the same evaluation with small supersonic wings similar to the [https://en.wikipedia.org/wiki/Pegasus_(rocket) Pegasus] ([http://www.aircraftdesign.com/pegasus_3view.gif wings image]). ==Evaluating lift and drag for a transonic/supersonic vehicle== The regular and accurate way to study aerodynamics is to use computational fluid dynamics (CFD). Some examples of that method can be seen [http://specificimpulses.blogspot.fr/search/label/Rocket%20Simulation here] for example. We will first look for approximations in standard conditions before trying this way, as we did for [[Heat transfer]], because CFD is quite complicated when you don't know how to use it, and CPU intensive. We consider here air-to-orbit rockets, so the subsonic part of the flight will be very short. We will ignore it for now, and directly skip to the transonic part. Our not-yet-published and approximative rocket flight and trajectory simulator informs us that the transonic regime lasts no more than 7 seconds if aerodynamic drag is ignored, with a release speed of Mach 0.9 and a thrust/weight ratio of 1.7. Most of the flight is thus supersonic and even hypersonic (Mach 5 should be reached at an altitude of 45km). ==Resources== It is quite easy to find information for model rockets with tail fins, mainly in subsonic flight. The best found so far is the '''OpenRocket technical documentation''' ([https://openrocket.sourceforge.net/techdoc.pdf pdf], 125 pages) from Sampo Niskanen, july 2011, based on his Master thesis. 8aecc791dfd164ec25b5fdedd6f27fdcde0f9e03 487 486 2013-01-01T02:38:41Z Vincent 1 another link: AeroFinSim and rocket category wikitext text/x-wiki =Rocket aerodynamics= We have seen with the [[Rocket:First_approximations|first approximation]] that the gravity drag is higher than expected for a single stage air-to-orbit configuration, when aerodynamic effects are ignored. The rocket engine has to be larger to compensate the gravity. On this page, we will evaluate how the aerodynamic effects can be used to compensate the gravity ([https://en.wikipedia.org/wiki/Lift_(force) lift]) without impacting the thrust too much ([https://en.wikipedia.org/wiki/Aerodynamic_drag drag]). This will be evaluated for a rocket without wings or fins at first, then we will do the same evaluation with small supersonic wings similar to the [https://en.wikipedia.org/wiki/Pegasus_(rocket) Pegasus] ([http://www.aircraftdesign.com/pegasus_3view.gif wings image]). ==Evaluating lift and drag for a transonic/supersonic vehicle== The regular and accurate way to study aerodynamics is to use computational fluid dynamics (CFD). Some examples of that method can be seen [http://specificimpulses.blogspot.fr/search/label/Rocket%20Simulation here] for example. We will first look for approximations in standard conditions before trying this way, as we did for [[Heat transfer]], because CFD is quite complicated when you don't know how to use it, and CPU intensive. We consider here air-to-orbit rockets, so the subsonic part of the flight will be very short. We will ignore it for now, and directly skip to the transonic part. Our not-yet-published and approximative rocket flight and trajectory simulator informs us that the transonic regime lasts no more than 7 seconds if aerodynamic drag is ignored, with a release speed of Mach 0.9 and a thrust/weight ratio of 1.7. Most of the flight is thus supersonic and even hypersonic (Mach 5 should be reached at an altitude of 45km). ==Resources== It is quite easy to find information for model rockets with tail fins, mainly in subsonic flight. The best found so far is the '''OpenRocket technical documentation''' ([https://openrocket.sourceforge.net/techdoc.pdf pdf], 125 pages) from Sampo Niskanen, july 2011, based on his Master thesis. The document is of very good quality and can be very useful even if it's not directly related to our flight conditions. In the final steps, if we use fins or small wings, we may use an aeroelastic analysis software like [http://www.aerorocket.com/finsim.html AeroFinSim] to design their physical properties. [[Category:Rocket]] 194692cd0f998ff1884d3d67d2846786130aff5e 488 487 2013-01-01T02:55:14Z Vincent 1 /* Resources */ fixing category wikitext text/x-wiki =Rocket aerodynamics= We have seen with the [[Rocket:First_approximations|first approximation]] that the gravity drag is higher than expected for a single stage air-to-orbit configuration, when aerodynamic effects are ignored. The rocket engine has to be larger to compensate the gravity. On this page, we will evaluate how the aerodynamic effects can be used to compensate the gravity ([https://en.wikipedia.org/wiki/Lift_(force) lift]) without impacting the thrust too much ([https://en.wikipedia.org/wiki/Aerodynamic_drag drag]). This will be evaluated for a rocket without wings or fins at first, then we will do the same evaluation with small supersonic wings similar to the [https://en.wikipedia.org/wiki/Pegasus_(rocket) Pegasus] ([http://www.aircraftdesign.com/pegasus_3view.gif wings image]). ==Evaluating lift and drag for a transonic/supersonic vehicle== The regular and accurate way to study aerodynamics is to use computational fluid dynamics (CFD). Some examples of that method can be seen [http://specificimpulses.blogspot.fr/search/label/Rocket%20Simulation here] for example. We will first look for approximations in standard conditions before trying this way, as we did for [[Heat transfer]], because CFD is quite complicated when you don't know how to use it, and CPU intensive. We consider here air-to-orbit rockets, so the subsonic part of the flight will be very short. We will ignore it for now, and directly skip to the transonic part. Our not-yet-published and approximative rocket flight and trajectory simulator informs us that the transonic regime lasts no more than 7 seconds if aerodynamic drag is ignored, with a release speed of Mach 0.9 and a thrust/weight ratio of 1.7. Most of the flight is thus supersonic and even hypersonic (Mach 5 should be reached at an altitude of 45km). ==Resources== It is quite easy to find information for model rockets with tail fins, mainly in subsonic flight. The best found so far is the '''OpenRocket technical documentation''' ([https://openrocket.sourceforge.net/techdoc.pdf pdf], 125 pages) from Sampo Niskanen, july 2011, based on his Master thesis. The document is of very good quality and can be very useful even if it's not directly related to our flight conditions. In the final steps, if we use fins or small wings, we may use an aeroelastic analysis software like [http://www.aerorocket.com/finsim.html AeroFinSim] to design their physical properties. [[Category:Rocket|Aerodynamics]] 3dbcac717903afb193384db9cc974b8987197a70 489 488 2013-01-03T02:51:33Z Vincent 1 /* Resources */ resources update wikitext text/x-wiki =Rocket aerodynamics= We have seen with the [[Rocket:First_approximations|first approximation]] that the gravity drag is higher than expected for a single stage air-to-orbit configuration, when aerodynamic effects are ignored. The rocket engine has to be larger to compensate the gravity. On this page, we will evaluate how the aerodynamic effects can be used to compensate the gravity ([https://en.wikipedia.org/wiki/Lift_(force) lift]) without impacting the thrust too much ([https://en.wikipedia.org/wiki/Aerodynamic_drag drag]). This will be evaluated for a rocket without wings or fins at first, then we will do the same evaluation with small supersonic wings similar to the [https://en.wikipedia.org/wiki/Pegasus_(rocket) Pegasus] ([http://www.aircraftdesign.com/pegasus_3view.gif wings image]). ==Evaluating lift and drag for a transonic/supersonic vehicle== The regular and accurate way to study aerodynamics is to use computational fluid dynamics (CFD). Some examples of that method can be seen [http://specificimpulses.blogspot.fr/search/label/Rocket%20Simulation here] for example. We will first look for approximations in standard conditions before trying this way, as we did for [[Heat transfer]], because CFD is quite complicated when you don't know how to use it, and CPU intensive. We consider here air-to-orbit rockets, so the subsonic part of the flight will be very short. We will ignore it for now, and directly skip to the transonic part. Our not-yet-published and approximative rocket flight and trajectory simulator informs us that the transonic regime lasts no more than 7 seconds if aerodynamic drag is ignored, with a release speed of Mach 0.9 and a thrust/weight ratio of 1.7. Most of the flight is thus supersonic and even hypersonic (Mach 5 should be reached at an altitude of 45km). ==Resources== It is quite easy to find information for model rockets with tail fins, mainly in subsonic flight. The best found so far is the '''OpenRocket technical documentation''' ([https://openrocket.sourceforge.net/techdoc.pdf pdf], 125 pages) from Sampo Niskanen, july 2011, based on his Master thesis. The document is of very good quality and can be very useful even if it's not directly related to our flight conditions. This [http://www.dept.aoe.vt.edu/~mason/Mason_f/ConfigAero.html Configuration Aerodynamics] class may be useful. ===Available Software=== [http://rasaero.com/ RASAero] created a nice tool for aerodynamics analysis running on Windows. It is free (costs no money): Rogers Aeroscience RASAero Aerodynamic Analysis and Flight Simulation Software. AeroRocket, a company of John Cipolla, has created several useful aerodynamics analysis tools, like [http://www.aerorocket.com/VisualCFD/Instructions.html VisualCFD] or [http://www.aerorocket.com/finsim.html AeroFinSim]. However, these tools are not free (it costs money) and also only work with Windows. [[Category:Rocket|Aerodynamics]] 50dd0542135a700e289c97bdb3b35dce3ac9dbd3 491 489 2013-01-05T23:06:48Z Vincent 1 max Q wikitext text/x-wiki =Rocket aerodynamics= We have seen with the [[Rocket:First_approximations|first approximation]] that the gravity drag is higher than expected for a single stage air-to-orbit configuration, when aerodynamic effects are ignored. The rocket engine has to be larger to compensate the gravity. On this page, we will evaluate how the aerodynamic effects can be used to compensate the gravity ([https://en.wikipedia.org/wiki/Lift_(force) lift]) without impacting the thrust too much ([https://en.wikipedia.org/wiki/Aerodynamic_drag drag]). This will be evaluated for a rocket without wings or fins at first, then we will do the same evaluation with small supersonic wings similar to the [https://en.wikipedia.org/wiki/Pegasus_(rocket) Pegasus] ([http://www.aircraftdesign.com/pegasus_3view.gif wings image]). ==Maximum dynamic pressure (max Q)== The main advantage of the air-to-orbit configuration is that the vehicle is not exposed to most of the atmosphere. We made a rocket flight trajectory simulator, and depending on the parameters of the vehicle, an approximation of the max Q can be calculated. For example, for a 1.7 thrust-to-weight ratio 652 kg vehicle given by our [[Rocket:First_approximations#Minimum_mass_evaluation|rocket mass program]], with a 30km altitude and 270m/s release speed, the '''max Q of 4664.36 Pa''' is reached after 54 seconds of flight, at a speed of Mach 3.40 and at an altitude of 34.8km. This program uses the [[Flight_at_high_altitude#Gas_properties_and_altitude|International Standard Atmosphere]] and a very rough estimate of the trajectory that needs to be hand-corrected for any parameter change, so it's not yet of publishable quality. ==Evaluating lift and drag for a transonic/supersonic vehicle== The regular and accurate way to study aerodynamics is to use computational fluid dynamics (CFD). Some examples of that method can be seen [http://specificimpulses.blogspot.fr/search/label/Rocket%20Simulation here] for example. We will first look for approximations in standard conditions before trying this way, as we did for [[Heat transfer]], because CFD is quite complicated when you don't know how to use it, and CPU intensive. We consider here air-to-orbit rockets, so the subsonic part of the flight will be very short. We will ignore it for now, and directly skip to the transonic part. Our not-yet-published and approximative rocket flight and trajectory simulator informs us that the transonic regime lasts no more than 7 seconds if aerodynamic drag is ignored, with a release speed of Mach 0.9 and a thrust/weight ratio of 1.7. Most of the flight is thus supersonic and even hypersonic (Mach 5 should be reached at an altitude of 45km). ==Resources== It is quite easy to find information for model rockets with tail fins, mainly in subsonic flight. The best found so far is the '''OpenRocket technical documentation''' ([https://openrocket.sourceforge.net/techdoc.pdf pdf], 125 pages) from Sampo Niskanen, july 2011, based on his Master thesis. The document is of very good quality and can be very useful even if it's not directly related to our flight conditions. This [http://www.dept.aoe.vt.edu/~mason/Mason_f/ConfigAero.html Configuration Aerodynamics] class may be useful. ===Available Software=== [http://rasaero.com/ RASAero] created a nice tool for aerodynamics analysis running on Windows. It is free (costs no money): Rogers Aeroscience RASAero Aerodynamic Analysis and Flight Simulation Software. AeroRocket, a company of John Cipolla, has created several useful aerodynamics analysis tools, like [http://www.aerorocket.com/VisualCFD/Instructions.html VisualCFD] or [http://www.aerorocket.com/finsim.html AeroFinSim]. However, these tools are not free (it costs money) and also only work with Windows. [[Category:Rocket|Aerodynamics]] b794a91aa8281e6111bd949177fd08882cc56095 492 491 2013-01-09T02:08:27Z Vincent 1 angle of attack wikitext text/x-wiki =Rocket aerodynamics= We have seen with the [[Rocket:First_approximations|first approximation]] that the gravity drag is higher than expected for a single stage air-to-orbit configuration, when aerodynamic effects are ignored. The rocket engine has to be larger to compensate the gravity. On this page, we will evaluate how the aerodynamic effects can be used to compensate the gravity ([https://en.wikipedia.org/wiki/Lift_(force) lift]) without impacting the thrust too much ([https://en.wikipedia.org/wiki/Aerodynamic_drag drag]). This will be evaluated for a rocket without wings or fins at first, then we will do the same evaluation with small supersonic wings similar to the [https://en.wikipedia.org/wiki/Pegasus_(rocket) Pegasus] ([http://www.aircraftdesign.com/pegasus_3view.gif wings image]). ==Maximum dynamic pressure (max Q)== The main advantage of the air-to-orbit configuration is that the vehicle is not exposed to most of the atmosphere. We made a rocket flight trajectory simulator, and depending on the parameters of the vehicle, an approximation of the max Q can be calculated. For example, for a 1.7 thrust-to-weight ratio 652 kg vehicle given by our [[Rocket:First_approximations#Minimum_mass_evaluation|rocket mass program]], with a 30km altitude and 270m/s release speed, the '''max Q of 4664.36 Pa''' is reached after 54 seconds of flight, at a speed of Mach 3.40 and at an altitude of 34.8km. This program uses the [[Flight_at_high_altitude#Gas_properties_and_altitude|International Standard Atmosphere]] and a very rough estimate of the trajectory that needs to be hand-corrected for any parameter change, so it's not yet of publishable quality. ==Evaluating lift and drag for a transonic/supersonic vehicle== The regular and accurate way to study aerodynamics is to use computational fluid dynamics (CFD). Some examples of that method can be seen [http://specificimpulses.blogspot.fr/search/label/Rocket%20Simulation here] for example. We will first look for approximations in standard conditions before trying this way, as we did for [[Heat transfer]], because CFD is quite complicated when you don't know how to use it, and CPU intensive. We consider here air-to-orbit rockets, so the subsonic part of the flight will be very short (after aircraft release). We will ignore it for now, and directly skip to the transonic part. Our not-yet-published and approximative rocket flight and trajectory simulator informs us that the transonic regime lasts no more than 7 seconds if aerodynamic drag is ignored, with a release speed of Mach 0.9 and a thrust/weight ratio of 1.7. Most of the flight is thus supersonic and even hypersonic (Mach 5 should be reached at an altitude of 45km). Another particularity of air-to-orbit vehicles is the high '''angle of attack'''. Indeed, contrary to balloon or ground launches, we already have an horizontal velocity on ignition. If the beginning of the trajectory is optimized to avoid staying in the low altitude atmosphere, gravity will need to be countered by pitching up aggressively. The body of the rocket will thus be at a high angle of attack, and same thing if supersonic wings are mounted on the rocket, they will not provide enough lift for the first few tens of seconds to change the velocity angle (real pitch). ==Resources== It is quite easy to find information for model rockets with tail fins, mainly in subsonic flight. The best found so far is the '''OpenRocket technical documentation''' ([https://openrocket.sourceforge.net/techdoc.pdf pdf], 125 pages) from Sampo Niskanen, july 2011, based on his Master thesis. The document is of very good quality and can be very useful even if it's not directly related to our flight conditions. This [http://www.dept.aoe.vt.edu/~mason/Mason_f/ConfigAero.html Configuration Aerodynamics] class may be useful. ===Available Software=== [http://rasaero.com/ RASAero] created a nice tool for aerodynamics analysis running on Windows. It is free (costs no money): Rogers Aeroscience RASAero Aerodynamic Analysis and Flight Simulation Software. AeroRocket, a company of John Cipolla, has created several useful aerodynamics analysis tools, like [http://www.aerorocket.com/VisualCFD/Instructions.html VisualCFD] or [http://www.aerorocket.com/finsim.html AeroFinSim]. However, these tools are not free (it costs money) and also only work with Windows. [[Category:Rocket|Aerodynamics]] 9867343153ecbd82029dcd51ac2bb1a7233f087f 493 492 2013-01-09T02:37:17Z Vincent 1 /* Resources */ RocketCalculator info wikitext text/x-wiki =Rocket aerodynamics= We have seen with the [[Rocket:First_approximations|first approximation]] that the gravity drag is higher than expected for a single stage air-to-orbit configuration, when aerodynamic effects are ignored. The rocket engine has to be larger to compensate the gravity. On this page, we will evaluate how the aerodynamic effects can be used to compensate the gravity ([https://en.wikipedia.org/wiki/Lift_(force) lift]) without impacting the thrust too much ([https://en.wikipedia.org/wiki/Aerodynamic_drag drag]). This will be evaluated for a rocket without wings or fins at first, then we will do the same evaluation with small supersonic wings similar to the [https://en.wikipedia.org/wiki/Pegasus_(rocket) Pegasus] ([http://www.aircraftdesign.com/pegasus_3view.gif wings image]). ==Maximum dynamic pressure (max Q)== The main advantage of the air-to-orbit configuration is that the vehicle is not exposed to most of the atmosphere. We made a rocket flight trajectory simulator, and depending on the parameters of the vehicle, an approximation of the max Q can be calculated. For example, for a 1.7 thrust-to-weight ratio 652 kg vehicle given by our [[Rocket:First_approximations#Minimum_mass_evaluation|rocket mass program]], with a 30km altitude and 270m/s release speed, the '''max Q of 4664.36 Pa''' is reached after 54 seconds of flight, at a speed of Mach 3.40 and at an altitude of 34.8km. This program uses the [[Flight_at_high_altitude#Gas_properties_and_altitude|International Standard Atmosphere]] and a very rough estimate of the trajectory that needs to be hand-corrected for any parameter change, so it's not yet of publishable quality. ==Evaluating lift and drag for a transonic/supersonic vehicle== The regular and accurate way to study aerodynamics is to use computational fluid dynamics (CFD). Some examples of that method can be seen [http://specificimpulses.blogspot.fr/search/label/Rocket%20Simulation here] for example. We will first look for approximations in standard conditions before trying this way, as we did for [[Heat transfer]], because CFD is quite complicated when you don't know how to use it, and CPU intensive. We consider here air-to-orbit rockets, so the subsonic part of the flight will be very short (after aircraft release). We will ignore it for now, and directly skip to the transonic part. Our not-yet-published and approximative rocket flight and trajectory simulator informs us that the transonic regime lasts no more than 7 seconds if aerodynamic drag is ignored, with a release speed of Mach 0.9 and a thrust/weight ratio of 1.7. Most of the flight is thus supersonic and even hypersonic (Mach 5 should be reached at an altitude of 45km). Another particularity of air-to-orbit vehicles is the high '''angle of attack'''. Indeed, contrary to balloon or ground launches, we already have an horizontal velocity on ignition. If the beginning of the trajectory is optimized to avoid staying in the low altitude atmosphere, gravity will need to be countered by pitching up aggressively. The body of the rocket will thus be at a high angle of attack, and same thing if supersonic wings are mounted on the rocket, they will not provide enough lift for the first few tens of seconds to change the velocity angle (real pitch). ==Resources== It is quite easy to find information for model rockets with tail fins, mainly in subsonic flight. The best found so far is the '''OpenRocket technical documentation''' ([https://openrocket.sourceforge.net/techdoc.pdf pdf], 125 pages) from Sampo Niskanen, july 2011, based on his Master thesis. The document is of very good quality and can be very useful even if it's not directly related to our flight conditions. This [http://www.dept.aoe.vt.edu/~mason/Mason_f/ConfigAero.html Configuration Aerodynamics] class may be useful. ===Available Software=== For speeds below Mach 3.0 and angle of attack below 25 degrees, RocketCalculator as described in the following paper can be used: Dahalan, Md. Nizam and Su, Vin Cent and Ammoo, Mohd. Shar (2009). '''Development of a computer program for rocket aerodynmic coefficients estimation'''. Jurnal Mekanikal, 28. pp.28-43 ([http://eprints.utm.my/21023/ link]). [http://rasaero.com/ RASAero] created a nice tool for aerodynamics analysis running on Windows. It is free (costs no money): Rogers Aeroscience RASAero Aerodynamic Analysis and Flight Simulation Software. AeroRocket, a company of John Cipolla, has created several useful aerodynamics analysis tools, like [http://www.aerorocket.com/VisualCFD/Instructions.html VisualCFD] or [http://www.aerorocket.com/finsim.html AeroFinSim]. However, these tools are not free (it costs money) and also only work with Windows. [[Category:Rocket|Aerodynamics]] 25349dac6c845f2833d5bdd4062cdd221fd00435 495 493 2013-02-12T01:23:15Z Vincent 1 /* Available Software */ update on rocketcalculator wikitext text/x-wiki =Rocket aerodynamics= We have seen with the [[Rocket:First_approximations|first approximation]] that the gravity drag is higher than expected for a single stage air-to-orbit configuration, when aerodynamic effects are ignored. The rocket engine has to be larger to compensate the gravity. On this page, we will evaluate how the aerodynamic effects can be used to compensate the gravity ([https://en.wikipedia.org/wiki/Lift_(force) lift]) without impacting the thrust too much ([https://en.wikipedia.org/wiki/Aerodynamic_drag drag]). This will be evaluated for a rocket without wings or fins at first, then we will do the same evaluation with small supersonic wings similar to the [https://en.wikipedia.org/wiki/Pegasus_(rocket) Pegasus] ([http://www.aircraftdesign.com/pegasus_3view.gif wings image]). ==Maximum dynamic pressure (max Q)== The main advantage of the air-to-orbit configuration is that the vehicle is not exposed to most of the atmosphere. We made a rocket flight trajectory simulator, and depending on the parameters of the vehicle, an approximation of the max Q can be calculated. For example, for a 1.7 thrust-to-weight ratio 652 kg vehicle given by our [[Rocket:First_approximations#Minimum_mass_evaluation|rocket mass program]], with a 30km altitude and 270m/s release speed, the '''max Q of 4664.36 Pa''' is reached after 54 seconds of flight, at a speed of Mach 3.40 and at an altitude of 34.8km. This program uses the [[Flight_at_high_altitude#Gas_properties_and_altitude|International Standard Atmosphere]] and a very rough estimate of the trajectory that needs to be hand-corrected for any parameter change, so it's not yet of publishable quality. ==Evaluating lift and drag for a transonic/supersonic vehicle== The regular and accurate way to study aerodynamics is to use computational fluid dynamics (CFD). Some examples of that method can be seen [http://specificimpulses.blogspot.fr/search/label/Rocket%20Simulation here] for example. We will first look for approximations in standard conditions before trying this way, as we did for [[Heat transfer]], because CFD is quite complicated when you don't know how to use it, and CPU intensive. We consider here air-to-orbit rockets, so the subsonic part of the flight will be very short (after aircraft release). We will ignore it for now, and directly skip to the transonic part. Our not-yet-published and approximative rocket flight and trajectory simulator informs us that the transonic regime lasts no more than 7 seconds if aerodynamic drag is ignored, with a release speed of Mach 0.9 and a thrust/weight ratio of 1.7. Most of the flight is thus supersonic and even hypersonic (Mach 5 should be reached at an altitude of 45km). Another particularity of air-to-orbit vehicles is the high '''angle of attack'''. Indeed, contrary to balloon or ground launches, we already have an horizontal velocity on ignition. If the beginning of the trajectory is optimized to avoid staying in the low altitude atmosphere, gravity will need to be countered by pitching up aggressively. The body of the rocket will thus be at a high angle of attack, and same thing if supersonic wings are mounted on the rocket, they will not provide enough lift for the first few tens of seconds to change the velocity angle (real pitch). ==Resources== It is quite easy to find information for model rockets with tail fins, mainly in subsonic flight. The best found so far is the '''OpenRocket technical documentation''' ([https://openrocket.sourceforge.net/techdoc.pdf pdf], 125 pages) from Sampo Niskanen, july 2011, based on his Master thesis. The document is of very good quality and can be very useful even if it's not directly related to our flight conditions. This [http://www.dept.aoe.vt.edu/~mason/Mason_f/ConfigAero.html Configuration Aerodynamics] class may be useful. ===Available Software=== [http://rasaero.com/ RASAero] created a nice tool for aerodynamics analysis running on Windows. It is free (costs no money): Rogers Aeroscience RASAero Aerodynamic Analysis and Flight Simulation Software. AeroRocket, a company of John Cipolla, has created several useful aerodynamics analysis tools, like [http://www.aerorocket.com/VisualCFD/Instructions.html VisualCFD] or [http://www.aerorocket.com/finsim.html AeroFinSim]. However, these tools are not free (it costs money) and also only work with Windows. For speeds below Mach 3.0 and angle of attack below 25 degrees, RocketCalculator as described in the following paper can be used: Dahalan, Md. Nizam and Su, Vin Cent and Ammoo, Mohd. Shar (2009). Development of a computer program for rocket aerodynmic coefficients estimation. Jurnal Mekanikal, 28. pp.28-43 ([http://eprints.utm.my/21023/ link]). '''''This program was requested several times using different communication ways, and no reply was received. Is it fake research?''''' [[Category:Rocket|Aerodynamics]] 1cb4154cd6aa7570ee058d76b869a175207ab5a1 496 495 2013-02-18T22:28:46Z Vincent 1 /* Resources */ link CFD book wikitext text/x-wiki =Rocket aerodynamics= We have seen with the [[Rocket:First_approximations|first approximation]] that the gravity drag is higher than expected for a single stage air-to-orbit configuration, when aerodynamic effects are ignored. The rocket engine has to be larger to compensate the gravity. On this page, we will evaluate how the aerodynamic effects can be used to compensate the gravity ([https://en.wikipedia.org/wiki/Lift_(force) lift]) without impacting the thrust too much ([https://en.wikipedia.org/wiki/Aerodynamic_drag drag]). This will be evaluated for a rocket without wings or fins at first, then we will do the same evaluation with small supersonic wings similar to the [https://en.wikipedia.org/wiki/Pegasus_(rocket) Pegasus] ([http://www.aircraftdesign.com/pegasus_3view.gif wings image]). ==Maximum dynamic pressure (max Q)== The main advantage of the air-to-orbit configuration is that the vehicle is not exposed to most of the atmosphere. We made a rocket flight trajectory simulator, and depending on the parameters of the vehicle, an approximation of the max Q can be calculated. For example, for a 1.7 thrust-to-weight ratio 652 kg vehicle given by our [[Rocket:First_approximations#Minimum_mass_evaluation|rocket mass program]], with a 30km altitude and 270m/s release speed, the '''max Q of 4664.36 Pa''' is reached after 54 seconds of flight, at a speed of Mach 3.40 and at an altitude of 34.8km. This program uses the [[Flight_at_high_altitude#Gas_properties_and_altitude|International Standard Atmosphere]] and a very rough estimate of the trajectory that needs to be hand-corrected for any parameter change, so it's not yet of publishable quality. ==Evaluating lift and drag for a transonic/supersonic vehicle== The regular and accurate way to study aerodynamics is to use computational fluid dynamics (CFD). Some examples of that method can be seen [http://specificimpulses.blogspot.fr/search/label/Rocket%20Simulation here] for example. We will first look for approximations in standard conditions before trying this way, as we did for [[Heat transfer]], because CFD is quite complicated when you don't know how to use it, and CPU intensive. We consider here air-to-orbit rockets, so the subsonic part of the flight will be very short (after aircraft release). We will ignore it for now, and directly skip to the transonic part. Our not-yet-published and approximative rocket flight and trajectory simulator informs us that the transonic regime lasts no more than 7 seconds if aerodynamic drag is ignored, with a release speed of Mach 0.9 and a thrust/weight ratio of 1.7. Most of the flight is thus supersonic and even hypersonic (Mach 5 should be reached at an altitude of 45km). Another particularity of air-to-orbit vehicles is the high '''angle of attack'''. Indeed, contrary to balloon or ground launches, we already have an horizontal velocity on ignition. If the beginning of the trajectory is optimized to avoid staying in the low altitude atmosphere, gravity will need to be countered by pitching up aggressively. The body of the rocket will thus be at a high angle of attack, and same thing if supersonic wings are mounted on the rocket, they will not provide enough lift for the first few tens of seconds to change the velocity angle (real pitch). ==Resources== It is quite easy to find information for model rockets with tail fins, mainly in subsonic flight. The best found so far is the '''OpenRocket technical documentation''' ([https://openrocket.sourceforge.net/techdoc.pdf pdf], 125 pages) from Sampo Niskanen, july 2011, based on his Master thesis. The document is of very good quality and can be very useful even if it's not directly related to our flight conditions. This [http://www.dept.aoe.vt.edu/~mason/Mason_f/ConfigAero.html Configuration Aerodynamics] class may be useful. If you're new to CFD, the following book is for you: John D. Anderson, Jr. '''Computational Fluid Dynamics: The Basics With Applications''', 1995. This is a beginners guide to CFD, quite outdated for direct software application but still well explained with lots of examples (poor quality nearly complete pdf [http://astronomy.nju.edu.cn/~chenpf/tmp/CFD.pdf here]). ===Available Software=== [http://rasaero.com/ RASAero] created a nice tool for aerodynamics analysis running on Windows. It is free (costs no money): Rogers Aeroscience RASAero Aerodynamic Analysis and Flight Simulation Software. AeroRocket, a company of John Cipolla, has created several useful aerodynamics analysis tools, like [http://www.aerorocket.com/VisualCFD/Instructions.html VisualCFD] or [http://www.aerorocket.com/finsim.html AeroFinSim]. However, these tools are not free (it costs money) and also only work with Windows. For speeds below Mach 3.0 and angle of attack below 25 degrees, RocketCalculator as described in the following paper can be used: Dahalan, Md. Nizam and Su, Vin Cent and Ammoo, Mohd. Shar (2009). Development of a computer program for rocket aerodynmic coefficients estimation. Jurnal Mekanikal, 28. pp.28-43 ([http://eprints.utm.my/21023/ link]). '''''This program was requested several times using different communication ways, and no reply was received. Is it fake research?''''' [[Category:Rocket|Aerodynamics]] bae75e6d3e00e0e8ccfa57305705ea8f01d8794a 499 496 2013-03-12T00:16:01Z Vincent 1 /* Rocket aerodynamics */ local backup for pegasus wings wikitext text/x-wiki =Rocket aerodynamics= We have seen with the [[Rocket:First_approximations|first approximation]] that the gravity drag is higher than expected for a single stage air-to-orbit configuration, when aerodynamic effects are ignored. The rocket engine has to be larger to compensate the gravity. On this page, we will evaluate how the aerodynamic effects can be used to compensate the gravity ([https://en.wikipedia.org/wiki/Lift_(force) lift]) without impacting the thrust too much ([https://en.wikipedia.org/wiki/Aerodynamic_drag drag]). This will be evaluated for a rocket without wings or fins at first, then we will do the same evaluation with small supersonic wings similar to the [https://en.wikipedia.org/wiki/Pegasus_(rocket) Pegasus] ([[:File:Pegasus_wings.gif|wings image]]). ==Maximum dynamic pressure (max Q)== The main advantage of the air-to-orbit configuration is that the vehicle is not exposed to most of the atmosphere. We made a rocket flight trajectory simulator, and depending on the parameters of the vehicle, an approximation of the max Q can be calculated. For example, for a 1.7 thrust-to-weight ratio 652 kg vehicle given by our [[Rocket:First_approximations#Minimum_mass_evaluation|rocket mass program]], with a 30km altitude and 270m/s release speed, the '''max Q of 4664.36 Pa''' is reached after 54 seconds of flight, at a speed of Mach 3.40 and at an altitude of 34.8km. This program uses the [[Flight_at_high_altitude#Gas_properties_and_altitude|International Standard Atmosphere]] and a very rough estimate of the trajectory that needs to be hand-corrected for any parameter change, so it's not yet of publishable quality. ==Evaluating lift and drag for a transonic/supersonic vehicle== The regular and accurate way to study aerodynamics is to use computational fluid dynamics (CFD). Some examples of that method can be seen [http://specificimpulses.blogspot.fr/search/label/Rocket%20Simulation here] for example. We will first look for approximations in standard conditions before trying this way, as we did for [[Heat transfer]], because CFD is quite complicated when you don't know how to use it, and CPU intensive. We consider here air-to-orbit rockets, so the subsonic part of the flight will be very short (after aircraft release). We will ignore it for now, and directly skip to the transonic part. Our not-yet-published and approximative rocket flight and trajectory simulator informs us that the transonic regime lasts no more than 7 seconds if aerodynamic drag is ignored, with a release speed of Mach 0.9 and a thrust/weight ratio of 1.7. Most of the flight is thus supersonic and even hypersonic (Mach 5 should be reached at an altitude of 45km). Another particularity of air-to-orbit vehicles is the high '''angle of attack'''. Indeed, contrary to balloon or ground launches, we already have an horizontal velocity on ignition. If the beginning of the trajectory is optimized to avoid staying in the low altitude atmosphere, gravity will need to be countered by pitching up aggressively. The body of the rocket will thus be at a high angle of attack, and same thing if supersonic wings are mounted on the rocket, they will not provide enough lift for the first few tens of seconds to change the velocity angle (real pitch). ==Resources== It is quite easy to find information for model rockets with tail fins, mainly in subsonic flight. The best found so far is the '''OpenRocket technical documentation''' ([https://openrocket.sourceforge.net/techdoc.pdf pdf], 125 pages) from Sampo Niskanen, july 2011, based on his Master thesis. The document is of very good quality and can be very useful even if it's not directly related to our flight conditions. This [http://www.dept.aoe.vt.edu/~mason/Mason_f/ConfigAero.html Configuration Aerodynamics] class may be useful. If you're new to CFD, the following book is for you: John D. Anderson, Jr. '''Computational Fluid Dynamics: The Basics With Applications''', 1995. This is a beginners guide to CFD, quite outdated for direct software application but still well explained with lots of examples (poor quality nearly complete pdf [http://astronomy.nju.edu.cn/~chenpf/tmp/CFD.pdf here]). ===Available Software=== [http://rasaero.com/ RASAero] created a nice tool for aerodynamics analysis running on Windows. It is free (costs no money): Rogers Aeroscience RASAero Aerodynamic Analysis and Flight Simulation Software. AeroRocket, a company of John Cipolla, has created several useful aerodynamics analysis tools, like [http://www.aerorocket.com/VisualCFD/Instructions.html VisualCFD] or [http://www.aerorocket.com/finsim.html AeroFinSim]. However, these tools are not free (it costs money) and also only work with Windows. For speeds below Mach 3.0 and angle of attack below 25 degrees, RocketCalculator as described in the following paper can be used: Dahalan, Md. Nizam and Su, Vin Cent and Ammoo, Mohd. Shar (2009). Development of a computer program for rocket aerodynmic coefficients estimation. Jurnal Mekanikal, 28. pp.28-43 ([http://eprints.utm.my/21023/ link]). '''''This program was requested several times using different communication ways, and no reply was received. Is it fake research?''''' [[Category:Rocket|Aerodynamics]] db23fd27a58f78ffefc215989e36a03cf01854e4 6 84 497 2013-03-12T00:14:03Z Vincent 1 3 views of Orbital Pegasus' wings wikitext text/x-wiki 3 views of Orbital Pegasus' wings 028d8fa33f94a1b896391d94d1ff7e7ef2a9b2a9 498 497 2013-03-12T00:15:39Z Vincent 1 source URL wikitext text/x-wiki 3 views of Orbital Pegasus' wings ([http://www.aircraftdesign.com/pegasus_3view.gif source]). 92d148c3ff9bb63d3a1a21da7eda84acd5415cef 0 85 500 2013-03-12T00:40:56Z Vincent 1 starting the CFD thing wikitext text/x-wiki This page has been created to help find resources on CFD, from a beginner level to achieve what we do in this project (nothing so far). =Introduction to Computational Fluid Dynamics= If you are new to CFD, a must read is this book, as recommended by [http://www.cfd-online.com/Books/show_book.php?book_id=3 cfd-online]: '''Computational Fluid Dynamics: The Basics with Applications''', by John David Anderson, 1995. Yes this book is nearly 20 years old, but is a very nice introduction to the topic, the theory doesn't age. If you want to see what CFD looks like on the theory and programming side, a brief overview of a basic method is presented in: [http://www.gputechconf.com/gtcnew/on-demand-gtc.php?sessionTopic=12&select=+#45 A Practical Introduction to Computational Fluid Dynamics on GPUs] from the GPU Tech Conference 2010 (direct link to [http://us.download.nvidia.com/downloads/GTC_Videos/flvs/2058_GTC2010.mp4 slides+voice mp4]). [http://www.cfd-online.com/Wiki/Main_Page CFD-Wiki], lots of useful pages and references. ffeabd518e23910b4418665f995d184586940f49 501 500 2013-03-12T01:39:59Z Vincent 1 /* Introduction to Computational Fluid Dynamics */ update on refs wikitext text/x-wiki This page has been created to help find resources on CFD, from a beginner level to achieve what we do in this project (nothing so far). =Introduction to Computational Fluid Dynamics= If you are new to CFD, a must read is this book, as recommended by [http://www.cfd-online.com/Books/show_book.php?book_id=3 cfd-online]: '''Computational Fluid Dynamics: The Basics with Applications''', by John David Anderson, 1995. Yes this book is nearly 20 years old, but is a very nice introduction to the topic, the theory doesn't age. A bad quality pdf is wandering on the Web. If you want to see what CFD looks like on the theory and programming side, a brief overview of a basic method is presented in: [http://www.gputechconf.com/gtcnew/on-demand-gtc.php?sessionTopic=12&select=+#45 A Practical Introduction to Computational Fluid Dynamics on GPUs] from the GPU Tech Conference 2010 (direct link to [http://us.download.nvidia.com/downloads/GTC_Videos/flvs/2058_GTC2010.mp4 slides+voice mp4] and [http://www.nvidia.com/content/GTC-2010/pdfs/2058_GTC2010.pdf presentation's pdf]). This presentation's main topic is using OpenCL to speed up CFD applications, so OpenCL is briefly presented too if you're interested. [http://www.cfd-online.com/Wiki/Main_Page CFD-Wiki], lots of useful pages and references. 62d9bf9fcd6454655bee491183bd076e3fe11aea 0 9 502 320 2013-03-14T02:12:01Z Vincent 1 /* Telemetry */ syrlinks link wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [https://www.armadeus.com/ Armadeus], based on an ARM (FreeScale) processor. Moreover, it [https://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [https://xenomai.org/ Xenomai] Linux-based RTOS. Armadeus board integration has a [[Armadeus|dedicated page]]. Other interesting embedded computer boards: the [https://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1] and the [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20]. They do not include a FPGA but are cheaper, the first has a great temperature range, and the second is more powerful and smaller. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [https://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. ===Telemetry/Communications=== It seems that the 900MHz version of the ZigBee communication standard is able to transmit at around 100kbps up to 10km. Taken from th ArduPilot page: Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [https://www.adafruit.com/products/126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [https://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [https://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. [http://www.syrlinks.com/en/products/cubesats.html Very High Data Rate Transmitter in X-Band for CubeSat]: This X-Band transmitter can transmit up to 13.3 GB per pass with a 5 meter station. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: quite common nowadays, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the theoretically computed strength the mechanical structure has to sustain. To chose a sensor, sparkfun wrote an [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and might be replaced by a camera sensor for low rotation rates. * Magnetometer (3D compass): lots of sensors exist too, for example the [https://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag] and [https://www.sparkfun.com/products/8128 SCP1000], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * Static pressure sensor, like [https://www.freescale.com/files/sensors/doc/data_sheet/MP3H6115A.pdf Freescale's MP3H6115A] and another for the extended range of high-altitude. The pressure sensor requires vents, or static ports, in the fuselage to exchange outside air with inside pressure chamber. Details about the size and number of these ports are explained [http://www.adeptrocketry.com/A1ds.htm here] for model rocketry. * Thermometer: for systems health monitoring, like engines temperature. * GPS if USAF and sensors allow it in flight altitude. * Camera: 8-bit data port if possible, like the TCM8230MD sensor. Some ARM processors (i.MX) feature the Camera/CMOS Sensor Interface (CSI) and hardware-accelerated processing or compression from this port. Horizon sensor is provided by a camera. * Pitot tubes even exist in stores (like [https://store.diydrones.com/Kit_MPXV7002DP_p/kt-mpxv7002dp-01.htm DIYDrones])! * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [https://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. 148655264de16804e9d7c30b87bd8e62e3dbc256 503 502 2013-03-14T02:13:48Z Vincent 1 /* Failsafe, mission abort */ link to mission page wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [https://www.armadeus.com/ Armadeus], based on an ARM (FreeScale) processor. Moreover, it [https://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [https://xenomai.org/ Xenomai] Linux-based RTOS. Armadeus board integration has a [[Armadeus|dedicated page]]. Other interesting embedded computer boards: the [https://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1] and the [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20]. They do not include a FPGA but are cheaper, the first has a great temperature range, and the second is more powerful and smaller. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [https://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. ===Telemetry/Communications=== It seems that the 900MHz version of the ZigBee communication standard is able to transmit at around 100kbps up to 10km. Taken from th ArduPilot page: Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [https://www.adafruit.com/products/126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [https://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [https://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. [http://www.syrlinks.com/en/products/cubesats.html Very High Data Rate Transmitter in X-Band for CubeSat]: This X-Band transmitter can transmit up to 13.3 GB per pass with a 5 meter station. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: quite common nowadays, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the theoretically computed strength the mechanical structure has to sustain. To chose a sensor, sparkfun wrote an [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and might be replaced by a camera sensor for low rotation rates. * Magnetometer (3D compass): lots of sensors exist too, for example the [https://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag] and [https://www.sparkfun.com/products/8128 SCP1000], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * Static pressure sensor, like [https://www.freescale.com/files/sensors/doc/data_sheet/MP3H6115A.pdf Freescale's MP3H6115A] and another for the extended range of high-altitude. The pressure sensor requires vents, or static ports, in the fuselage to exchange outside air with inside pressure chamber. Details about the size and number of these ports are explained [http://www.adeptrocketry.com/A1ds.htm here] for model rocketry. * Thermometer: for systems health monitoring, like engines temperature. * GPS if USAF and sensors allow it in flight altitude. * Camera: 8-bit data port if possible, like the TCM8230MD sensor. Some ARM processors (i.MX) feature the Camera/CMOS Sensor Interface (CSI) and hardware-accelerated processing or compression from this port. Horizon sensor is provided by a camera. * Pitot tubes even exist in stores (like [https://store.diydrones.com/Kit_MPXV7002DP_p/kt-mpxv7002dp-01.htm DIYDrones])! * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [https://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. Options are explored on the [[Aircraft_Mission#Staging_and_risk_evaluation|mission page]]. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. 9a6867ece9bb6fdaa9f391b7940078cb3013f757 507 503 2013-08-07T13:23:21Z Vincent 1 /* High processing power */ beagle board link wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [https://www.armadeus.com/ Armadeus], based on an ARM (FreeScale) processor. Moreover, it [https://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [https://xenomai.org/ Xenomai] Linux-based RTOS. Armadeus board integration has a [[Armadeus|dedicated page]]. Other interesting embedded computer boards: the [https://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1], [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20] and the [http://beagleboard.org/ BeagleBoard]. They do not include a FPGA but are cheaper, the first has a great temperature range, the second is more powerful and smaller, and the third is cheap, open and has a large community. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [https://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. ===Telemetry/Communications=== It seems that the 900MHz version of the ZigBee communication standard is able to transmit at around 100kbps up to 10km. Taken from th ArduPilot page: Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [https://www.adafruit.com/products/126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [https://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [https://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. [http://www.syrlinks.com/en/products/cubesats.html Very High Data Rate Transmitter in X-Band for CubeSat]: This X-Band transmitter can transmit up to 13.3 GB per pass with a 5 meter station. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: quite common nowadays, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the theoretically computed strength the mechanical structure has to sustain. To chose a sensor, sparkfun wrote an [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and might be replaced by a camera sensor for low rotation rates. * Magnetometer (3D compass): lots of sensors exist too, for example the [https://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag] and [https://www.sparkfun.com/products/8128 SCP1000], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * Static pressure sensor, like [https://www.freescale.com/files/sensors/doc/data_sheet/MP3H6115A.pdf Freescale's MP3H6115A] and another for the extended range of high-altitude. The pressure sensor requires vents, or static ports, in the fuselage to exchange outside air with inside pressure chamber. Details about the size and number of these ports are explained [http://www.adeptrocketry.com/A1ds.htm here] for model rocketry. * Thermometer: for systems health monitoring, like engines temperature. * GPS if USAF and sensors allow it in flight altitude. * Camera: 8-bit data port if possible, like the TCM8230MD sensor. Some ARM processors (i.MX) feature the Camera/CMOS Sensor Interface (CSI) and hardware-accelerated processing or compression from this port. Horizon sensor is provided by a camera. * Pitot tubes even exist in stores (like [https://store.diydrones.com/Kit_MPXV7002DP_p/kt-mpxv7002dp-01.htm DIYDrones])! * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [https://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. Options are explored on the [[Aircraft_Mission#Staging_and_risk_evaluation|mission page]]. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. 4f5b653bb0e23abd2c642d2bf94d5f2f617201ec 508 507 2013-08-07T13:26:36Z Vincent 1 /* Low processing power */ Gravity board link wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [https://www.armadeus.com/ Armadeus], based on an ARM (FreeScale) processor. Moreover, it [https://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [https://xenomai.org/ Xenomai] Linux-based RTOS. Armadeus board integration has a [[Armadeus|dedicated page]]. Other interesting embedded computer boards: the [https://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1], [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20] and the [http://beagleboard.org/ BeagleBoard]. They do not include a FPGA but are cheaper, the first has a great temperature range, the second is more powerful and smaller, and the third is cheap, open and has a large community. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [https://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. Other Arduino-based initiatives have appeared. For example, the [https://github.com/zortness/rocket-mega-shield Rocket Mega shield] features a fast GPS chip, XBee telemetry, MicroSD card, pressure/altitude sensor, gyroscope and accelerometers. A few other microcontroller boards have recently received some attention, like the [http://www.solarsystemexpress.com/gravity-development-board.html GravityBoard]. It is more powerful than Arduinos, features high-power outputs, uses space-tolerant hardware. ===Telemetry/Communications=== It seems that the 900MHz version of the ZigBee communication standard is able to transmit at around 100kbps up to 10km. Taken from th ArduPilot page: Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [https://www.adafruit.com/products/126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [https://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [https://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. [http://www.syrlinks.com/en/products/cubesats.html Very High Data Rate Transmitter in X-Band for CubeSat]: This X-Band transmitter can transmit up to 13.3 GB per pass with a 5 meter station. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: quite common nowadays, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the theoretically computed strength the mechanical structure has to sustain. To chose a sensor, sparkfun wrote an [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and might be replaced by a camera sensor for low rotation rates. * Magnetometer (3D compass): lots of sensors exist too, for example the [https://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag] and [https://www.sparkfun.com/products/8128 SCP1000], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * Static pressure sensor, like [https://www.freescale.com/files/sensors/doc/data_sheet/MP3H6115A.pdf Freescale's MP3H6115A] and another for the extended range of high-altitude. The pressure sensor requires vents, or static ports, in the fuselage to exchange outside air with inside pressure chamber. Details about the size and number of these ports are explained [http://www.adeptrocketry.com/A1ds.htm here] for model rocketry. * Thermometer: for systems health monitoring, like engines temperature. * GPS if USAF and sensors allow it in flight altitude. * Camera: 8-bit data port if possible, like the TCM8230MD sensor. Some ARM processors (i.MX) feature the Camera/CMOS Sensor Interface (CSI) and hardware-accelerated processing or compression from this port. Horizon sensor is provided by a camera. * Pitot tubes even exist in stores (like [https://store.diydrones.com/Kit_MPXV7002DP_p/kt-mpxv7002dp-01.htm DIYDrones])! * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [https://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. Options are explored on the [[Aircraft_Mission#Staging_and_risk_evaluation|mission page]]. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. ec35277d33e4ff7b6f05e65349b0b98a0464f61e 509 508 2013-08-07T13:46:58Z Vincent 1 /* Sensors */ A small sensors update wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [https://www.armadeus.com/ Armadeus], based on an ARM (FreeScale) processor. Moreover, it [https://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [https://xenomai.org/ Xenomai] Linux-based RTOS. Armadeus board integration has a [[Armadeus|dedicated page]]. Other interesting embedded computer boards: the [https://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1], [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20] and the [http://beagleboard.org/ BeagleBoard]. They do not include a FPGA but are cheaper, the first has a great temperature range, the second is more powerful and smaller, and the third is cheap, open and has a large community. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [https://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. Other Arduino-based initiatives have appeared. For example, the [https://github.com/zortness/rocket-mega-shield Rocket Mega shield] features a fast GPS chip, XBee telemetry, MicroSD card, pressure/altitude sensor, gyroscope and accelerometers. A few other microcontroller boards have recently received some attention, like the [http://www.solarsystemexpress.com/gravity-development-board.html GravityBoard]. It is more powerful than Arduinos, features high-power outputs, uses space-tolerant hardware. ===Telemetry/Communications=== It seems that the 900MHz version of the ZigBee communication standard is able to transmit at around 100kbps up to 10km. Taken from th ArduPilot page: Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [https://www.adafruit.com/products/126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [https://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [https://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. [http://www.syrlinks.com/en/products/cubesats.html Very High Data Rate Transmitter in X-Band for CubeSat]: This X-Band transmitter can transmit up to 13.3 GB per pass with a 5 meter station. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: nowadays common, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the stress the structure has to sustain. To chose a sensor, sparkfun wrote an [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and might be replaced by a camera sensor for low rotation rates. * Magnetometer (3D compass): lots of sensors exist too, for example the [https://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag] and [https://www.sparkfun.com/products/8128 SCP1000], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * Static pressure sensor, like [https://www.freescale.com/files/sensors/doc/data_sheet/MP3H6115A.pdf Freescale's MP3H6115A] and another for the extended range of high-altitude. Small and very sensitive sensors are appearing, like the [http://www.digikey.com/product-detail/en/BMP085/828-1005-1-ND/1987010 BMP085]. The pressure sensor requires vents, or static ports, in the fuselage to exchange outside air with inside pressure chamber. Details about the size and number of these ports are explained [http://www.adeptrocketry.com/A1ds.htm here] for model rocketry. * Thermometer: for systems health monitoring, like engines temperature. * GPS: sensors have an altitude limitation due to USAF export regulation, but some unlimited chips can easily be found. 10Hz refresh rate can be obtained for relatively cheap ($30), like the [http://www.adafruit.com/products/790 MTK3339]. * Camera: 8-bit data port if possible, like the TCM8230MD sensor. Some ARM processors (i.MX) feature the Camera/CMOS Sensor Interface (CSI) and hardware-accelerated processing or compression from this port. Horizon sensor is provided by a camera. * Pitot tubes even exist in stores (like [https://store.diydrones.com/Kit_MPXV7002DP_p/kt-mpxv7002dp-01.htm DIYDrones])! * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [https://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. Options are explored on the [[Aircraft_Mission#Staging_and_risk_evaluation|mission page]]. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. d9b2b280ccfee8c459419e9f3576f185247f9633 511 509 2013-08-14T17:04:13Z Vincent 1 /* Telemetry/Communications */ xbee info and ardupilot link wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [https://www.armadeus.com/ Armadeus], based on an ARM (FreeScale) processor. Moreover, it [https://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [https://xenomai.org/ Xenomai] Linux-based RTOS. Armadeus board integration has a [[Armadeus|dedicated page]]. Other interesting embedded computer boards: the [https://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1], [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20] and the [http://beagleboard.org/ BeagleBoard]. They do not include a FPGA but are cheaper, the first has a great temperature range, the second is more powerful and smaller, and the third is cheap, open and has a large community. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [https://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. Other Arduino-based initiatives have appeared. For example, the [https://github.com/zortness/rocket-mega-shield Rocket Mega shield] features a fast GPS chip, XBee telemetry, MicroSD card, pressure/altitude sensor, gyroscope and accelerometers. A few other microcontroller boards have recently received some attention, like the [http://www.solarsystemexpress.com/gravity-development-board.html GravityBoard]. It is more powerful than Arduinos, features high-power outputs, uses space-tolerant hardware. ===Telemetry/Communications=== From what is explained on SparkFun's [https://www.sparkfun.com/pages/xbee_guide XBee guide], the 900MHz low data rate (10kbps) version of XBee is able to transmit up to a range of 24km (15 miles), when using a high gain antenna. Also see the [http://code.google.com/p/ardupilot-mega/wiki/Telem ArduPilot telemetry page]. Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [https://www.adafruit.com/products/126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [https://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [https://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. [http://www.syrlinks.com/en/products/cubesats.html Very High Data Rate Transmitter in X-Band for CubeSat]: This X-Band transmitter can transmit up to 13.3 GB per pass with a 5 meter station. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: nowadays common, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the stress the structure has to sustain. To chose a sensor, sparkfun wrote an [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and might be replaced by a camera sensor for low rotation rates. * Magnetometer (3D compass): lots of sensors exist too, for example the [https://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag] and [https://www.sparkfun.com/products/8128 SCP1000], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * Static pressure sensor, like [https://www.freescale.com/files/sensors/doc/data_sheet/MP3H6115A.pdf Freescale's MP3H6115A] and another for the extended range of high-altitude. Small and very sensitive sensors are appearing, like the [http://www.digikey.com/product-detail/en/BMP085/828-1005-1-ND/1987010 BMP085]. The pressure sensor requires vents, or static ports, in the fuselage to exchange outside air with inside pressure chamber. Details about the size and number of these ports are explained [http://www.adeptrocketry.com/A1ds.htm here] for model rocketry. * Thermometer: for systems health monitoring, like engines temperature. * GPS: sensors have an altitude limitation due to USAF export regulation, but some unlimited chips can easily be found. 10Hz refresh rate can be obtained for relatively cheap ($30), like the [http://www.adafruit.com/products/790 MTK3339]. * Camera: 8-bit data port if possible, like the TCM8230MD sensor. Some ARM processors (i.MX) feature the Camera/CMOS Sensor Interface (CSI) and hardware-accelerated processing or compression from this port. Horizon sensor is provided by a camera. * Pitot tubes even exist in stores (like [https://store.diydrones.com/Kit_MPXV7002DP_p/kt-mpxv7002dp-01.htm DIYDrones])! * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [https://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. Options are explored on the [[Aircraft_Mission#Staging_and_risk_evaluation|mission page]]. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. 72862c55eeefd80b89798965e51475d19e35bf0c 512 511 2013-08-21T16:57:23Z Vincent 1 /* Telemetry/Communications */ radio antics link wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a very important part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of constraints: power consumption, size, weight, operating temperature... Outcomes are seen in processing power, memory space, connectivity (I/O ports), battery life, and mechanical design. An embedded control computer has to have a low latency to process data from attitude sensors and command actuators. Realtime computing must be achieved through a hard-realtime operating system, or without using an operating system if you have only one process. ==Hardware== It's hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and their processing. Beyond that, we only need to get the command control, mission planning, and telemetry, that don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curve for positioning, it will require a lot of processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [https://www.armadeus.com/ Armadeus], based on an ARM (FreeScale) processor. Moreover, it [https://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [https://xenomai.org/ Xenomai] Linux-based RTOS. Armadeus board integration has a [[Armadeus|dedicated page]]. Other interesting embedded computer boards: the [https://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1], [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20] and the [http://beagleboard.org/ BeagleBoard]. They do not include a FPGA but are cheaper, the first has a great temperature range, the second is more powerful and smaller, and the third is cheap, open and has a large community. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [https://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. Other Arduino-based initiatives have appeared. For example, the [https://github.com/zortness/rocket-mega-shield Rocket Mega shield] features a fast GPS chip, XBee telemetry, MicroSD card, pressure/altitude sensor, gyroscope and accelerometers. A few other microcontroller boards have recently received some attention, like the [http://www.solarsystemexpress.com/gravity-development-board.html GravityBoard]. It is more powerful than Arduinos, features high-power outputs, uses space-tolerant hardware. ===Telemetry/Communications=== From what is explained on SparkFun's [https://www.sparkfun.com/pages/xbee_guide XBee guide], the 900MHz low data rate (10kbps) version of XBee is able to transmit up to a range of 24km (15 miles), when using a high gain antenna. Also see the [http://code.google.com/p/ardupilot-mega/wiki/Telem ArduPilot telemetry page]. Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [https://www.adafruit.com/products/126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [https://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [https://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. [http://www.syrlinks.com/en/products/cubesats.html Very High Data Rate Transmitter in X-Band for CubeSat]: This X-Band transmitter can transmit up to 13.3 GB per pass with a 5 meter station. "I have since discovered that the transmitters in these balloons are only 10mW, but because of the line of sight they can manage several hundred miles", in 434MHz band, from [http://nerdsville.blogspot.co.uk/2013/04/99-red-balloons-well-actually-3-high.html Radio Antics]. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: nowadays common, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the stress the structure has to sustain. To chose a sensor, sparkfun wrote an [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and might be replaced by a camera sensor for low rotation rates. * Magnetometer (3D compass): lots of sensors exist too, for example the [https://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag] and [https://www.sparkfun.com/products/8128 SCP1000], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * Static pressure sensor, like [https://www.freescale.com/files/sensors/doc/data_sheet/MP3H6115A.pdf Freescale's MP3H6115A] and another for the extended range of high-altitude. Small and very sensitive sensors are appearing, like the [http://www.digikey.com/product-detail/en/BMP085/828-1005-1-ND/1987010 BMP085]. The pressure sensor requires vents, or static ports, in the fuselage to exchange outside air with inside pressure chamber. Details about the size and number of these ports are explained [http://www.adeptrocketry.com/A1ds.htm here] for model rocketry. * Thermometer: for systems health monitoring, like engines temperature. * GPS: sensors have an altitude limitation due to USAF export regulation, but some unlimited chips can easily be found. 10Hz refresh rate can be obtained for relatively cheap ($30), like the [http://www.adafruit.com/products/790 MTK3339]. * Camera: 8-bit data port if possible, like the TCM8230MD sensor. Some ARM processors (i.MX) feature the Camera/CMOS Sensor Interface (CSI) and hardware-accelerated processing or compression from this port. Horizon sensor is provided by a camera. * Pitot tubes even exist in stores (like [https://store.diydrones.com/Kit_MPXV7002DP_p/kt-mpxv7002dp-01.htm DIYDrones])! * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [https://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. Options are explored on the [[Aircraft_Mission#Staging_and_risk_evaluation|mission page]]. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. 83ccce815b7352db2a5da85ff587b32bb7086517 513 512 2013-08-23T01:54:16Z Vincent 1 constraints and environment of space computers. Adding the link to telemetry page too. wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a major part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of '''constraints''': power consumption, size, weight, heat and operating temperature... On the other hand, it must have a low latency to process the attitude control loop (read data from attitude sensors and command the actuators). Realtime computing is a standard for microcontroller-based systems, which do not feature an operating system overlay, but it can be more complicated to have deadline guarantees when an OS is used, on PC-based boards. In that case, a hard-realtime operating system must be used, unless you are willing to replace it with your sole process and manage all I/O yourself. The high altitude and space environments are also raise the problem of '''radiation'''. For electronic systems, there are mainly two issues: bit flip or component burning because of [https://en.wikipedia.org/wiki/Radiation_hardening#Major_radiation_damage_sources radiation] and sparking because of potential differences on sun-facing and opposite external faces of a satellite's structure. The space radiation environment is thoroughly described on this [http://www.diyspaceexploration.com/space-radiation-effects-on-electronic-components-in-low-earth-orbit/ DYI Space Exploration page], based on a NASA report. [https://en.wikipedia.org/wiki/Radiation_hardening#Radiation-hardening_techniques Many techniques] exist against radiation, both physical with radiation-hardened electronics and with software features like redundancy. ==Hardware== It's obviously hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and the processing power they require. Beyond that, we only need to get the command control, mission planning and telemetry, which don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curvature for attitude sensing, it will require a lot more processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [https://www.armadeus.com/ Armadeus], based on an ARM (FreeScale) processor. Moreover, it [https://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [https://xenomai.org/ Xenomai] Linux-based RTOS. Armadeus board integration has a [[Armadeus|dedicated page]]. Other interesting embedded computer boards: the [https://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1], [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20] and the [http://beagleboard.org/ BeagleBoard]. They do not include a FPGA but are cheaper, the first has a great temperature range, the second is more powerful and smaller, and the third is cheap, open and has a large community. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [https://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. Other Arduino-based initiatives have appeared. For example, the [https://github.com/zortness/rocket-mega-shield Rocket Mega shield] features a fast GPS chip, XBee telemetry, MicroSD card, pressure/altitude sensor, gyroscope and accelerometers. A few other microcontroller boards have recently received some attention, like the [http://www.solarsystemexpress.com/gravity-development-board.html GravityBoard]. It is more powerful than Arduinos, features high-power outputs, uses space-tolerant hardware. ===Telemetry/Communications=== ''Main article: [[Telemetry]]'' From what is explained on SparkFun's [https://www.sparkfun.com/pages/xbee_guide XBee guide], the 900MHz low data rate (10kbps) version of XBee is able to transmit up to a range of 24km (15 miles), when using a high gain antenna. Also see the [http://code.google.com/p/ardupilot-mega/wiki/Telem ArduPilot telemetry page]. Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [https://www.adafruit.com/products/126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [https://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [https://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. [http://www.syrlinks.com/en/products/cubesats.html Very High Data Rate Transmitter in X-Band for CubeSat]: This X-Band transmitter can transmit up to 13.3 GB per pass with a 5 meter station. "I have since discovered that the transmitters in these balloons are only 10mW, but because of the line of sight they can manage several hundred miles", in 434MHz band, from [http://nerdsville.blogspot.co.uk/2013/04/99-red-balloons-well-actually-3-high.html Radio Antics]. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: nowadays common, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the stress the structure has to sustain. To chose a sensor, sparkfun wrote an [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and might be replaced by a camera sensor for low rotation rates. * Magnetometer (3D compass): lots of sensors exist too, for example the [https://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag] and [https://www.sparkfun.com/products/8128 SCP1000], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * Static pressure sensor, like [https://www.freescale.com/files/sensors/doc/data_sheet/MP3H6115A.pdf Freescale's MP3H6115A] and another for the extended range of high-altitude. Small and very sensitive sensors are appearing, like the [http://www.digikey.com/product-detail/en/BMP085/828-1005-1-ND/1987010 BMP085]. The pressure sensor requires vents, or static ports, in the fuselage to exchange outside air with inside pressure chamber. Details about the size and number of these ports are explained [http://www.adeptrocketry.com/A1ds.htm here] for model rocketry. * Thermometer: for systems health monitoring, like engines temperature. * GPS: sensors have an altitude limitation due to USAF export regulation, but some unlimited chips can easily be found. 10Hz refresh rate can be obtained for relatively cheap ($30), like the [http://www.adafruit.com/products/790 MTK3339]. * Camera: 8-bit data port if possible, like the TCM8230MD sensor. Some ARM processors (i.MX) feature the Camera/CMOS Sensor Interface (CSI) and hardware-accelerated processing or compression from this port. Horizon sensor is provided by a camera. * Pitot tubes even exist in stores (like [https://store.diydrones.com/Kit_MPXV7002DP_p/kt-mpxv7002dp-01.htm DIYDrones])! * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [https://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. Options are explored on the [[Aircraft_Mission#Staging_and_risk_evaluation|mission page]]. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. c573cb00e2d83a72f063ebf34ca44d065b3e4803 525 513 2013-09-26T14:01:11Z Vincent 1 /* Sensors */ pressure sensor for rocket engine thrust calculation wikitext text/x-wiki =Embedded computer: attitude and mission control, telemetry= The embedded computer is a major part of a launcher, because of the development and testing time it requires, and because a simple unforeseen case can lead the whole operation to failure. The embedded computing world undergoes lots of '''constraints''': power consumption, size, weight, heat and operating temperature... On the other hand, it must have a low latency to process the attitude control loop (read data from attitude sensors and command the actuators). Realtime computing is a standard for microcontroller-based systems, which do not feature an operating system overlay, but it can be more complicated to have deadline guarantees when an OS is used, on PC-based boards. In that case, a hard-realtime operating system must be used, unless you are willing to replace it with your sole process and manage all I/O yourself. The high altitude and space environments are also raise the problem of '''radiation'''. For electronic systems, there are mainly two issues: bit flip or component burning because of [https://en.wikipedia.org/wiki/Radiation_hardening#Major_radiation_damage_sources radiation] and sparking because of potential differences on sun-facing and opposite external faces of a satellite's structure. The space radiation environment is thoroughly described on this [http://www.diyspaceexploration.com/space-radiation-effects-on-electronic-components-in-low-earth-orbit/ DYI Space Exploration page], based on a NASA report. [https://en.wikipedia.org/wiki/Radiation_hardening#Radiation-hardening_techniques Many techniques] exist against radiation, both physical with radiation-hardened electronics and with software features like redundancy. ==Hardware== It's obviously hard to have a low cost, small form factor, with high processing power. But do we really need high power? That depends on what sensors are used, and the processing power they require. Beyond that, we only need to get the command control, mission planning and telemetry, which don't require a high power. Since we are limited by cost, we won't be able to get high quality sensors, or highly integrated sensors like an inertial sensor, but rather accelerometers, and digital gyroscopes. Their data will need to be processed, but that still does not require a lot of power. If we use a video camera however, to track the sun and the Earth's curvature for attitude sensing, it will require a lot more processing power. ===High processing power=== An alternative to pure processing power by a CPU exists: digital signal processors (DSPs), and since they are also very expensive, we can use FPGAs to program them. An FPGA (Field-Programmable Gate Array) is an electronic chip with a matrix of gates that can be programmed in order to specialize it to a specific information processing. It then acts as a hardware processing based on a software definition of the processing, offloading the CPU. Information about FPGAs can be found at [http://www.fpga4fun.com fpga4fun.com]. And it happens that there is an embedded microprocessor board that includes a FPGA and multiple I/Os, and a quite faire processing power: the [https://www.armadeus.com/ Armadeus], based on an ARM (FreeScale) processor. Moreover, it [https://www.armadeus.com/wiki/index.php?title=Xenomai supports] the free (GPL) [https://xenomai.org/ Xenomai] Linux-based RTOS. Armadeus board integration has a [[Armadeus|dedicated page]]. Other interesting embedded computer boards: the [https://shop.trenz-electronic.de/catalog/product_info.php?cPath=26_55_116&products_id=541 Eddy-CPU v2.1], [http://www.taskit.de/en/products/portuxg20/index.htm Portux G20] and the [http://beagleboard.org/ BeagleBoard]. They do not include a FPGA but are cheaper, the first has a great temperature range, the second is more powerful and smaller, and the third is cheap, open and has a large community. ===Low processing power=== If video is not used as a sensor, microcontrollers may be able to handle some sensors and actuators, at least for aircraft control. The [https://diydrones.com/profiles/blogs/ardupilot-main-page ArduPilot] is a good example of open project trying to achieve that. Other Arduino-based initiatives have appeared. For example, the [https://github.com/zortness/rocket-mega-shield Rocket Mega shield] features a fast GPS chip, XBee telemetry, MicroSD card, pressure/altitude sensor, gyroscope and accelerometers. A few other microcontroller boards have recently received some attention, like the [http://www.solarsystemexpress.com/gravity-development-board.html GravityBoard]. It is more powerful than Arduinos, features high-power outputs, uses space-tolerant hardware. ===Telemetry/Communications=== ''Main article: [[Telemetry]]'' From what is explained on SparkFun's [https://www.sparkfun.com/pages/xbee_guide XBee guide], the 900MHz low data rate (10kbps) version of XBee is able to transmit up to a range of 24km (15 miles), when using a high gain antenna. Also see the [http://code.google.com/p/ardupilot-mega/wiki/Telem ArduPilot telemetry page]. Two Xbee modules for wireless telemetry: [https://www.sparkfun.com/commerce/product_info.php?products_id=9097 This one] with [https://www.adafruit.com/products/126 this adapter] in the air and [https://www.sparkfun.com/commerce/product_info.php?products_id=9099 this one] with [https://www.sparkfun.com/commerce/product_info.php?products_id=9143 this antenna] and [https://www.sparkfun.com/commerce/product_info.php?products_id=8687 this adapter board]. [http://www.syrlinks.com/en/products/cubesats.html Very High Data Rate Transmitter in X-Band for CubeSat]: This X-Band transmitter can transmit up to 13.3 GB per pass with a 5 meter station. "I have since discovered that the transmitters in these balloons are only 10mW, but because of the line of sight they can manage several hundred miles", in 434MHz band, from [http://nerdsville.blogspot.co.uk/2013/04/99-red-balloons-well-actually-3-high.html Radio Antics]. ===Sensors=== Before creating a new dedicated [[Sensors]] page because it takes too much space here, here is a list of sensors that can or should be used: * Accelerometers: nowadays common, accelerometers allow attitude sensing, together with gyroscopes or/and magnetometers. For our project, a single-axis accelerometer can be used to detect free-fall created by separation of plane and rocket, and to sense the roll movement of the rocket, at least for the first part of the flight, since Earth gravity will be more or less sensed depending on the roll. A second accelerometer could be used for thrust confirmation, collinear to the length of the rocket. It would also be a nice telemetry feature, and provide a feedback on the stress the structure has to sustain. To chose a sensor, sparkfun wrote an [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167 accelerometer tutorial]. * Gyroscopes: they can obviously be helpful on attitude sensing, for yaw roll and pitch of the plane, and thus for the control command. To chose a sensor, sparkfun wrote a [https://www.sparkfun.com/commerce/tutorial_info.php?tutorials_id=167&sipp=1&page=2 gyroscope tutorial]. However, those sensors can be relatively expensive for a decent precision, and might be replaced by a camera sensor for low rotation rates. * Magnetometer (3D compass): lots of sensors exist too, for example the [https://www.sparkfun.com/commerce/product_info.php?products_id=244 MicroMag] and [https://www.sparkfun.com/products/8128 SCP1000], but are quite expensive. Knowing where is the North of Earth can be very useful, in order to corroborate information from the camera or other sensors, and add some precision to the orbital injection parameters. * Static pressure sensor, like [https://www.freescale.com/files/sensors/doc/data_sheet/MP3H6115A.pdf Freescale's MP3H6115A] and another for the extended range of high-altitude. Small and very sensitive sensors are appearing, like the [http://www.digikey.com/product-detail/en/BMP085/828-1005-1-ND/1987010 BMP085]. The pressure sensor requires vents, or static ports, in the fuselage to exchange outside air with inside pressure chamber. Details about the size and number of these ports are explained [http://www.adeptrocketry.com/A1ds.htm here] for model rocketry. * Thermometer: for systems health monitoring, like engines temperature. * GPS: sensors have an altitude limitation due to USAF export regulation, but some unlimited chips can easily be found. 10Hz refresh rate can be obtained for relatively cheap ($30), like the [http://www.adafruit.com/products/790 MTK3339]. * Camera: 8-bit data port if possible, like the TCM8230MD sensor. Some ARM processors (i.MX) feature the Camera/CMOS Sensor Interface (CSI) and hardware-accelerated processing or compression from this port. Horizon sensor is provided by a camera. * Pitot tubes even exist in stores (like [https://store.diydrones.com/Kit_MPXV7002DP_p/kt-mpxv7002dp-01.htm DIYDrones])! * Fuel gauge or low level indicator and thus end of mission, orbital injection parameters freezing and stating. For the rocket in particular, a pressure sensor is needed for the rocket engine. It is both used to verify that everything is working fine, and it allows the thrust to be computed in flight. As said on [[Resources#Web_pages|arocket]]: ''I'm convinced that if you have a good chamber pressure transducer, know the throat diameter, and have good estimates of thrust coefficient, it's probably more accurate than directly measuring thrust from a load cell''. ==Software== First thing about software is always thinking about the model of the application, meaning how will it be conceived or organized. Several layers are generally seen in softwares: * Real application: mission ** Keep track of the status in the mission ** Send orders (commands) to the control layer * Control system ** Sensors and actuator communication and processing ** Control loop from sensors to actuators regarding to commands * Operating system * Hardware ===Mission: the launch program=== We need to chose a way to express and manage the mission. It is defined by actions to trigger when some conditions are met, like "when altitude is 60km, proceed to staging", or "at T+7s, begin roll program". ===Control=== The [https://en.wikipedia.org/wiki/Control_system control loop]'s purpose is to ensure that the vehicle is in a state consistent with the state expected by the mission. It controls attitude (roll, pitch, yaw) of the vehicle in order to make it fit with the expected attitude. In our case, roll is not really a concern for the rocket, since the satellite does not carry important science payload that has to be pointed in a particular direction. For the aircraft, on the other side, it is very important. A control loop is decomposed like that: ''picture'' Sensors information is collected and processed. Actuator commands are processed from both sensor data and expected-to-be-reached sensor data (nominal flight pattern). This loop has to be processed several times per second, with a highly accurate timing. Indeed, sensor processing, for example accelerometer data, has to be integrated to know the speed and the position of the vehicle. If time shifts randomly, calculated speed will not be correct, leading to false actuation command. With no luck, and we have to assume that it is the case, that creates real attitude error while it was not previously bad. If error is too important on pitch for example, it can lead to catastrophic structural damage at such high speeds. Hard realtime operating systems (RTOS) guarantee that the time between expected processing time and actual processing time (the system's latency) is bounded by a very low maximum value. ===Failsafe, mission abort=== In case something goes wrong, for example and engine failure, or structural failure, if it can be detected by sensors, the systems will have to go into a failsafe mode - basically shutting down everything that can explode and try to return to ground in the minimum of different parts. Options are explored on the [[Aircraft_Mission#Staging_and_risk_evaluation|mission page]]. In some cases, the mission will need to be aborted from ground, because no sensor was available for a specific task, or because of a programming error. An upload communication would then be needed, allowing the system to cease its mission, and try to go back in one piece. 953c535389b5ba51495b209b30aa96f1f1d029ff 0 1 504 494 2013-03-22T02:08:28Z Vincent 1 /* News */ CFD 2013Q1 wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''March 2013:''''' Since december, learning CFD has been the main activity, and it will probably remain so next months. [[CFD:Introduction|A page]] has been created to give CFD beginners some interesting links. CFD has a steep learning curve, but learning how to use it will have huge benefits for the project in the long term: * evaluate the lift and drag associated with supersonic wings/fins on the aircraft-launched rocket will enable us to refine the rocket mass ([[File:Rocket_mass.c]]) model * hopefully have a theoretical validation of our high-altitude turbofan [[Turbofan:Alternative_Designs|alternative design]] * simulate different wing profiles at high-altitude subsonic conditions for the carrier aircraft * evaluate aircraft and engine capability on lift-off and early flight conditions while they are both tailored for high altitude flight * refine the heat transfer approximations ([[File:Heat_transfer_to_rocket_tank.c]]) for cryogenics tank vaporization, used in the rocket mass model for tank dimensioning, with a better climb profile input too. '''''December 2012:''''' A simple rocket trajectory model has been made in order to evaluate the trajectory of an aircraft- or balloon-released-rocket. It appears that the Delta V taken for granted for gravity drag for these rockets, around 800 m/s, is quite erroneous, or not possible with the expected overall thrust-to-weight ratio. Without taking into account the aerodynamic effects like lift and drag, for a ratio of around 1.5, the Delta V for gravity drag is at least 1300 m/s. To reduce it, the ratio should be higher, like 3.0, in that case it may be possible to have only 800 m/s Delta V for gravity, but the mass of the engine would be much higher. See [[Rocket:First_approximations#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|a solution]] to this issue. The program evaluating the mass of the rocket has been refined (v3), and the forgotten pipe linking the upper tank to the engine adds approximately 150 kg of wet mass too ([[File:Rocket_mass.c]]). Incoming works will study aerodynamic lift and drag of a simple rocket and of a rocket with small supersonic wings to better evaluate the gravity drag for our mass evaluation. '''''November 2012:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 150 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] have been created to illustrate the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. The program ([[File:Rocket_mass.c]]) modelling the mass of rockets is being refined and the cryogenic propellant vaporization has been evaluated by another program ([[File:Heat_transfer_to_rocket_tank.c]]). Both programs are freely available. '''''August 2012:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012:''''' Study of basic aerodynamics is under way. More man power is expected in April. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. See [http://gravityloss.wordpress.com/2008/04/21/air-breathers-advantage this article] for more information on the differences of rocket and aircraft propulsion efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several N-Prize teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. In [http://www.youtube.com/watch?v=esgc5W_Ufng this video] (SpaceX guys), here captured at SpaceUP, they don't even predict the use of attitude control outside the atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell ([http://www.jpaerospace.com/ JP Aeroospace]) is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) equivalent to the empty weight of the plane and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Small gas turbine engines exist however, even in an axial design, see [http://www.bladonjets.com/technology/gas-turbines/ Bladon Jets]. Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]], Category page: [[:Category:Rocket|list of pages]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost, but also regarding their efficiency, the Isp they can produce. This is [[Rocket:First_approximations#Effects_of_parameter_changes|especially important]] in an SSTO design. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Turbopumps are used on full-scale commercial rockets, but are very complex to build and design. Xcor has created and demonstrated since 2003 a [http://www.xcor.com/products/pumps/ piston pump] for LOX, which is now used on a 1,500 lb-thrust (6.6 kN) LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory is closely tied to the [[Flight_at_high_altitude#Approaches_overview|initial release parameters]], the [[Rocket:First_approximations|flight parameters]] such as thrust and aerodynamics, and [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control], which depends on sensors and actuators: '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. 01f6fcff14bcd3a9b3de98bdbd93198e85f6f002 533 504 2013-10-01T00:05:52Z Vincent 1 /* The aircraft */ adding price induction turbofan link wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''March 2013:''''' Since december, learning CFD has been the main activity, and it will probably remain so next months. [[CFD:Introduction|A page]] has been created to give CFD beginners some interesting links. CFD has a steep learning curve, but learning how to use it will have huge benefits for the project in the long term: * evaluate the lift and drag associated with supersonic wings/fins on the aircraft-launched rocket will enable us to refine the rocket mass ([[File:Rocket_mass.c]]) model * hopefully have a theoretical validation of our high-altitude turbofan [[Turbofan:Alternative_Designs|alternative design]] * simulate different wing profiles at high-altitude subsonic conditions for the carrier aircraft * evaluate aircraft and engine capability on lift-off and early flight conditions while they are both tailored for high altitude flight * refine the heat transfer approximations ([[File:Heat_transfer_to_rocket_tank.c]]) for cryogenics tank vaporization, used in the rocket mass model for tank dimensioning, with a better climb profile input too. '''''December 2012:''''' A simple rocket trajectory model has been made in order to evaluate the trajectory of an aircraft- or balloon-released-rocket. It appears that the Delta V taken for granted for gravity drag for these rockets, around 800 m/s, is quite erroneous, or not possible with the expected overall thrust-to-weight ratio. Without taking into account the aerodynamic effects like lift and drag, for a ratio of around 1.5, the Delta V for gravity drag is at least 1300 m/s. To reduce it, the ratio should be higher, like 3.0, in that case it may be possible to have only 800 m/s Delta V for gravity, but the mass of the engine would be much higher. See [[Rocket:First_approximations#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|a solution]] to this issue. The program evaluating the mass of the rocket has been refined (v3), and the forgotten pipe linking the upper tank to the engine adds approximately 150 kg of wet mass too ([[File:Rocket_mass.c]]). Incoming works will study aerodynamic lift and drag of a simple rocket and of a rocket with small supersonic wings to better evaluate the gravity drag for our mass evaluation. '''''November 2012:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 150 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] have been created to illustrate the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. The program ([[File:Rocket_mass.c]]) modelling the mass of rockets is being refined and the cryogenic propellant vaporization has been evaluated by another program ([[File:Heat_transfer_to_rocket_tank.c]]). Both programs are freely available. '''''August 2012:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012:''''' Study of basic aerodynamics is under way. More man power is expected in April. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. See [http://gravityloss.wordpress.com/2008/04/21/air-breathers-advantage this article] for more information on the differences of rocket and aircraft propulsion efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several N-Prize teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. In [http://www.youtube.com/watch?v=esgc5W_Ufng this video] (SpaceX guys), here captured at SpaceUP, they don't even predict the use of attitude control outside the atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell ([http://www.jpaerospace.com/ JP Aeroospace]) is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) equivalent to the empty weight of the plane and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Small gas turbine engines exist however, even in an axial design, see [http://www.bladonjets.com/technology/gas-turbines/ Bladon Jets] or [http://www.price-induction.com/en/dgen/caracteristiques/ Price Induction] for a commercial turbofan for very small jets (nice picture of the fan [http://www.price-induction.com/site_media/images/dgen-net/technologies/optimisation_de_masse_grand.jpg here]). Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]], Category page: [[:Category:Rocket|list of pages]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost, but also regarding their efficiency, the Isp they can produce. This is [[Rocket:First_approximations#Effects_of_parameter_changes|especially important]] in an SSTO design. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Turbopumps are used on full-scale commercial rockets, but are very complex to build and design. Xcor has created and demonstrated since 2003 a [http://www.xcor.com/products/pumps/ piston pump] for LOX, which is now used on a 1,500 lb-thrust (6.6 kN) LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory is closely tied to the [[Flight_at_high_altitude#Approaches_overview|initial release parameters]], the [[Rocket:First_approximations|flight parameters]] such as thrust and aerodynamics, and [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control], which depends on sensors and actuators: '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. 93a978a6a44ac63681c644ccb8c36b86851bd50d 542 533 2013-12-26T01:10:25Z Vincent 1 /* News */ telemetry wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''December 2013:''''' CFD has been stalled for a few months. In parallel, work on [[Telemetry|RF telemetry]] has started. First prototype will be tested in a high-altitude balloon during the beginning of 2014, in order to validate the RF module and the ground receiver. '''''March 2013:''''' Since December, learning CFD has been the main activity, and it will probably remain so next months. [[CFD:Introduction|A page]] has been created to give CFD beginners some interesting links. CFD has a steep learning curve, but learning how to use it will have huge benefits for the project in the long term: * evaluate the lift and drag associated with supersonic wings/fins on the aircraft-launched rocket will enable us to refine the rocket mass ([[File:Rocket_mass.c]]) model * hopefully have a theoretical validation of our high-altitude turbofan [[Turbofan:Alternative_Designs|alternative design]] * simulate different wing profiles at high-altitude subsonic conditions for the carrier aircraft * evaluate aircraft and engine capability on lift-off and early flight conditions while they are both tailored for high altitude flight * refine the heat transfer approximations ([[File:Heat_transfer_to_rocket_tank.c]]) for cryogenics tank vaporization, used in the rocket mass model for tank dimensioning, with a better climb profile input too. '''''December 2012:''''' A simple rocket trajectory model has been made in order to evaluate the trajectory of an aircraft- or balloon-released-rocket. It appears that the Delta V taken for granted for gravity drag for these rockets, around 800 m/s, is quite erroneous, or not possible with the expected overall thrust-to-weight ratio. Without taking into account the aerodynamic effects like lift and drag, for a ratio of around 1.5, the Delta V for gravity drag is at least 1300 m/s. To reduce it, the ratio should be higher, like 3.0, in that case it may be possible to have only 800 m/s Delta V for gravity, but the mass of the engine would be much higher. See [[Rocket:First_approximations#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|a solution]] to this issue. The program evaluating the mass of the rocket has been refined (v3), and the forgotten pipe linking the upper tank to the engine adds approximately 150 kg of wet mass too ([[File:Rocket_mass.c]]). Incoming works will study aerodynamic lift and drag of a simple rocket and of a rocket with small supersonic wings to better evaluate the gravity drag for our mass evaluation. '''''November 2012:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 150 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] have been created to illustrate the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. The program ([[File:Rocket_mass.c]]) modelling the mass of rockets is being refined and the cryogenic propellant vaporization has been evaluated by another program ([[File:Heat_transfer_to_rocket_tank.c]]). Both programs are freely available. '''''August 2012:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012:''''' Study of basic aerodynamics is under way. More man power is expected in April. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. See [http://gravityloss.wordpress.com/2008/04/21/air-breathers-advantage this article] for more information on the differences of rocket and aircraft propulsion efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several N-Prize teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. In [http://www.youtube.com/watch?v=esgc5W_Ufng this video] (SpaceX guys), here captured at SpaceUP, they don't even predict the use of attitude control outside the atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell ([http://www.jpaerospace.com/ JP Aeroospace]) is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) equivalent to the empty weight of the plane and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Small gas turbine engines exist however, even in an axial design, see [http://www.bladonjets.com/technology/gas-turbines/ Bladon Jets] or [http://www.price-induction.com/en/dgen/caracteristiques/ Price Induction] for a commercial turbofan for very small jets (nice picture of the fan [http://www.price-induction.com/site_media/images/dgen-net/technologies/optimisation_de_masse_grand.jpg here]). Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]], Category page: [[:Category:Rocket|list of pages]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost, but also regarding their efficiency, the Isp they can produce. This is [[Rocket:First_approximations#Effects_of_parameter_changes|especially important]] in an SSTO design. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Turbopumps are used on full-scale commercial rockets, but are very complex to build and design. Xcor has created and demonstrated since 2003 a [http://www.xcor.com/products/pumps/ piston pump] for LOX, which is now used on a 1,500 lb-thrust (6.6 kN) LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory is closely tied to the [[Flight_at_high_altitude#Approaches_overview|initial release parameters]], the [[Rocket:First_approximations|flight parameters]] such as thrust and aerodynamics, and [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control], which depends on sensors and actuators: '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. 2b0bb77266fedae19b74461f6ee6a5c90195fb66 0 16 505 439 2013-04-04T01:05:04Z Vincent 1 /* Web pages */ adding Ed Williams and unreasonablerocket wikitext text/x-wiki =Resources= This page gathers available documentation on the numerous subjects linked the project, including rocket science, turbofans, aerodynamics, amateur rocketry and SSTO. An [[Aero formulas|internal page]] has been created containing useful formulas related to thermodynamics and aerodynamics. Another page contains all [[heat transfer]]-related material. ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/ amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * An important nasaspaceflight [http://forum.nasaspaceflight.com/index.php?topic=2847.0 forum thread] where ideas on micro-rocketry to orbit are discussed and shared. * A kind of spin-off of the above thread is the [http://orbitalaspirations.blogspot.com/ Orbital Aspirations weblog]. It was recently created by Ed LeBouthillier and is already filled with lots of information on scaling down rocket equations to micro-rocketry, SSTO, reference papers, news of the domain and so on. * Another blog talking about SSTO and air-to-orbit: [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Selenian Boondocks]. * Another blog is dedicated to SSTO: [http://exoscientist.blogspot.fr/ Polymath/Exoscientist]. * A scientist blog with some occasional rocketry: [http://www.gravityloss.com/ Gravity Loss]. * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. * [http://williams.best.vwh.net/index.html Ed Williams' Aviation page] Flight dynamics and formulas about aviation in general. A good start to understand flight controls. Link found on this [http://unreasonablerocket.blogspot.fr/2013/04/teaching-computer-to-fly.html unreasonable rocket] blog entry about UAV. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] by Klaus Hünecke, 1997. * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ===Other: Thermodynamics, Aerodynamics, general physics=== * [http://www.desktop.aero/appliedaero/preface/welcome.html Applied Aerodynamics: A Digital Textbook] by Desktop Aeronautics, Inc. 2007. * [http://www.rshanthini.com/ThermoBook.htm Thermodynamics for Beginners, with worked examples] by R. Shanthini. 2006. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. 493035e383baa09d6369df340e50e4fae8d555fd 526 505 2013-09-26T14:13:50Z Vincent 1 adding similar projects wikitext text/x-wiki =Resources= This page gathers available documentation on the numerous subjects linked the project, including rocket science, turbofans, aerodynamics, amateur rocketry and SSTO. An [[Aero formulas|internal page]] has been created containing useful formulas related to thermodynamics and aerodynamics. Another page contains all [[heat transfer]]-related material. ==Similar projects== * [http://www.opennautics.com/ Opennautics] (European project) develops a hybrid engine for an amateur rocket, all open source. * [http://www.wikisat.org/?page_id=2 WikiSat] (European project, mostly Spanish) work on pico satellites and their launcher, mainly a rocket launched from a balloon (a rockoon). ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/ amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * An important nasaspaceflight [http://forum.nasaspaceflight.com/index.php?topic=2847.0 forum thread] where ideas on micro-rocketry to orbit are discussed and shared. * A kind of spin-off of the above thread is the [http://orbitalaspirations.blogspot.com/ Orbital Aspirations weblog]. It was recently created by Ed LeBouthillier and is already filled with lots of information on scaling down rocket equations to micro-rocketry, SSTO, reference papers, news of the domain and so on. * Another blog talking about SSTO and air-to-orbit: [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Selenian Boondocks]. * Another blog is dedicated to SSTO: [http://exoscientist.blogspot.fr/ Polymath/Exoscientist]. * A scientist blog with some occasional rocketry: [http://www.gravityloss.com/ Gravity Loss]. * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. * [http://williams.best.vwh.net/index.html Ed Williams' Aviation page] Flight dynamics and formulas about aviation in general. A good start to understand flight controls. Link found on this [http://unreasonablerocket.blogspot.fr/2013/04/teaching-computer-to-fly.html unreasonable rocket] blog entry about UAV. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] by Klaus Hünecke, 1997. * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ===Other: Thermodynamics, Aerodynamics, general physics=== * [http://www.desktop.aero/appliedaero/preface/welcome.html Applied Aerodynamics: A Digital Textbook] by Desktop Aeronautics, Inc. 2007. * [http://www.rshanthini.com/ThermoBook.htm Thermodynamics for Beginners, with worked examples] by R. Shanthini. 2006. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. 257a69b258a6c0dd055b83ff612acdcd62e06910 543 526 2013-12-26T01:16:02Z Vincent 1 /* Similar projects */ SRRG wikitext text/x-wiki =Resources= This page gathers available documentation on the numerous subjects linked the project, including rocket science, turbofans, aerodynamics, amateur rocketry and SSTO. An [[Aero formulas|internal page]] has been created containing useful formulas related to thermodynamics and aerodynamics. Another page contains all [[heat transfer]]-related material. ==Similar projects== * [http://www.opennautics.com/ Opennautics] (European project) develops a hybrid engine for an amateur rocket, all open source. * [http://www.wikisat.org/?page_id=2 WikiSat] (European project, mostly Spanish) work on pico satellites and their launcher, mainly a rocket launched from a balloon (a rockoon). * [http://swedenrocketresearchgroup.blogspot.fr/ Sweden Rocket Research Group], also an amateur project developing rocket engines and accessories, currently a hybrid engine. ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/ amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * An important nasaspaceflight [http://forum.nasaspaceflight.com/index.php?topic=2847.0 forum thread] where ideas on micro-rocketry to orbit are discussed and shared. * A kind of spin-off of the above thread is the [http://orbitalaspirations.blogspot.com/ Orbital Aspirations weblog]. It was recently created by Ed LeBouthillier and is already filled with lots of information on scaling down rocket equations to micro-rocketry, SSTO, reference papers, news of the domain and so on. * Another blog talking about SSTO and air-to-orbit: [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Selenian Boondocks]. * Another blog is dedicated to SSTO: [http://exoscientist.blogspot.fr/ Polymath/Exoscientist]. * A scientist blog with some occasional rocketry: [http://www.gravityloss.com/ Gravity Loss]. * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. * [http://williams.best.vwh.net/index.html Ed Williams' Aviation page] Flight dynamics and formulas about aviation in general. A good start to understand flight controls. Link found on this [http://unreasonablerocket.blogspot.fr/2013/04/teaching-computer-to-fly.html unreasonable rocket] blog entry about UAV. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] by Klaus Hünecke, 1997. * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ===Other: Thermodynamics, Aerodynamics, general physics=== * [http://www.desktop.aero/appliedaero/preface/welcome.html Applied Aerodynamics: A Digital Textbook] by Desktop Aeronautics, Inc. 2007. * [http://www.rshanthini.com/ThermoBook.htm Thermodynamics for Beginners, with worked examples] by R. Shanthini. 2006. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. e81366811feabd9873126c90f2f474ca5b43ad3c 547 543 2013-12-31T14:04:43Z Vincent 1 /* Similar projects */ srrg youtube link wikitext text/x-wiki =Resources= This page gathers available documentation on the numerous subjects linked the project, including rocket science, turbofans, aerodynamics, amateur rocketry and SSTO. An [[Aero formulas|internal page]] has been created containing useful formulas related to thermodynamics and aerodynamics. Another page contains all [[heat transfer]]-related material. ==Similar projects== * [http://www.opennautics.com/ Opennautics] (European project) develops a hybrid engine for an amateur rocket, all open source. * [http://www.wikisat.org/?page_id=2 WikiSat] (European project, mostly Spanish) work on pico satellites and their launcher, mainly a rocket launched from a balloon (a rockoon). * [http://swedenrocketresearchgroup.blogspot.fr/ Sweden Rocket Research Group], also an amateur project developing rocket engines and accessories, currently a hybrid engine. Link to their [https://www.youtube.com/user/borgaren84/videos?shelf_index=0&sort=dd&tag_id=&view=0 youtube videos]. ==Web pages== A more complete and multi-domain list of links is available on the [http://www.arocketry.net/ amateur rocketry website], as well as on the ARocket mailing list, subscription available on this same site. Most useful links are below: * An important nasaspaceflight [http://forum.nasaspaceflight.com/index.php?topic=2847.0 forum thread] where ideas on micro-rocketry to orbit are discussed and shared. * A kind of spin-off of the above thread is the [http://orbitalaspirations.blogspot.com/ Orbital Aspirations weblog]. It was recently created by Ed LeBouthillier and is already filled with lots of information on scaling down rocket equations to micro-rocketry, SSTO, reference papers, news of the domain and so on. * Another blog talking about SSTO and air-to-orbit: [http://selenianboondocks.blogspot.fr/2007/01/orbital-access-methodologies-part-i-air.html Selenian Boondocks]. * Another blog is dedicated to SSTO: [http://exoscientist.blogspot.fr/ Polymath/Exoscientist]. * A scientist blog with some occasional rocketry: [http://www.gravityloss.com/ Gravity Loss]. * Robert A. Braeunig's [http://www.braeunig.us/space/ website] on rocket and space technology, including a nice forum. * [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm Nozzle design]. * [http://www.xcor.com/products/pumps/ XCOR cryogenic piston pumps] (for LOX) on [http://www.xcor.com/products/engines/4K5_LOX-Kerosene_rocket_engine.html XR-4K5], a 1,800 lbf LOX/kerosene Engine. * [http://www.mentallandscape.com/S_R7.htm History of the R7] (soyuz rocket) and rocket engines issues prior to it. * [http://williams.best.vwh.net/index.html Ed Williams' Aviation page] Flight dynamics and formulas about aviation in general. A good start to understand flight controls. Link found on this [http://unreasonablerocket.blogspot.fr/2013/04/teaching-computer-to-fly.html unreasonable rocket] blog entry about UAV. ==Lectures== * [http://ocw.mit.edu/ans7870/16/16.unified/propulsionS04/UnifiedPropulsion1/UnifiedPropulsion1.htm Unified Propulsion MIT lecture]. Lots of interesting stuff, especially in page 9 "Energy Exchange with Moving Blades". * [http://mit.edu/16.unified/www/FALL/thermodynamics/notes/notes.html Thermodynamics and Propulsion MIT lecture]. Lots of interesting stuff too, especially the third part on propulsion. ==Books (online links)== ===Multi-domain=== * [http://books.google.com/books?id=jM4yNV5xTscC Aerothermodynamics of gas turbine and rocket propulsion] by Gordon C. Oates. 1997. ===Rocket engines=== * [http://nprize.mine.nu/~vinvin/rocket_book/ How to design, build and test small liquid-fuel rocket engines] by Rocketlab / China lake, Calif. 1967. ''Local copy of the full book''. Backup is [[:Image:Rocket_book.tar.gz|here]]. * [http://books.google.com/books?id=LQbDOxg3XZcC Rocket propulsion elements] by George Paul Sutton and Oscar Biblarz. 7th edition, 2001. ''Very complete.'' * [http://books.google.com/books?id=TKdIbLX51NQC Modern engineering for design of liquid-propellant rocket engines] by Dieter K. Huzel, David H. Huang and Harry Arbit. 1992. * [http://www.spl.ch/publication/sp125.html The Design of Liquid Propellant Rockets] (full book) 2nd edition by Huzel and Huang, 1971. ====Specific topics of rocket egines:==== * [http://books.google.com/books?id=1OC8zeol7uMC Cryogenic engineering] by Thomas M. Flynn. 2005. * [http://books.google.com/books?id=sobvSF82RVAC Liquid rocket engine combustion instability] by Vigor Yang and William E. Anderson. 1995. * [http://books.google.com/books?id=0HWotm1k40QC Liquid rocket thrust chambers: aspects of modeling, analysis, and design] by Vigor Yang. 2004. ===Turbines and turbofans=== * [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation] by Klaus Hünecke, 1997. * [http://books.google.com/books?id=V0SnFt8JGokC The history of North American small gas turbine aircraft engines] by Richard A. Leyes and William A. Fleming. 1999. * [http://books.google.com/books?id=dZkDQgAACAAJ Turbofan] by Frederic P Miller, Agnes F Vandome and John McBrewster. 2009. * [http://books.google.com/books?id=yy2YoIKDC3gC Jet propulsion:] a simple guide to the aerodynamics and thermodynamic design and performance of jet engines, by N. A. Cumpsty. 2003. * [http://books.google.com/books?id=ummg5F227WoC Gas turbine theory] by H. I. H. Saravanamuttoo, Gordon Frederick Crichton Rogers and Henry Cohen. 2001. * [http://books.google.com/books?id=zPAFHyHctRUC Gas turbine engineering handbook] by Meherwan P. Boyce. 2002. ===Other: Thermodynamics, Aerodynamics, general physics=== * [http://www.desktop.aero/appliedaero/preface/welcome.html Applied Aerodynamics: A Digital Textbook] by Desktop Aeronautics, Inc. 2007. * [http://www.rshanthini.com/ThermoBook.htm Thermodynamics for Beginners, with worked examples] by R. Shanthini. 2006. ==Videos (youtube links)== * [http://www.youtube.com/watch?v=esgc5W_Ufng SSTO presentation] at SpaceUp conference. * [http://www.youtube.com/watch?v=vN3_Wkyl5PQ How are made turbine blades] from Discovey channel "How it's made" show. * [http://www.youtube.com/watch?v=G7LQuVfDwFQ Using propellers to get into space] by [http://www.jpaerospace.com/ John Powell] in SpaceVidCast. * [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk Video documentaries on turbine engines], from a turbine renovator in Canada, probably the best resource on the Web for seing what's inside real engines. 3fa46b1ce9b41808787772d6791edeb308e99ba8 0 75 506 481 2013-06-26T21:16:11Z Vincent 1 /* References */ FUBAR labs wiki on heat wikitext text/x-wiki [https://en.wikipedia.org/wiki/Heat_transfer Heat transfer] is a very complicated process involving many parameters and conditions. This page plays the role of introduction to heat transfer and a documentation on the methods used to calculate the heat transfers applied to our vehicle. =Heat transfers= A heat transfer is basically an energy transfer that can produce work or a change of temperature difference between the interacting parts. It can take three forms: conductive, convective and radiative. In the case of a heat transfer due to the atmosphere around an aircraft, the three mechanisms are effective. This page currently focuses on convective heat transfer, which implies conductive heat transfer. Radiative transfer is probably negligible before the two others in our particular case study. ==Introduction to heat convection and conduction== Heat '''convection''' occurs when there is a fluid flowing around a control volume at a temperature different than the control volume's. The flow can be either ''natural'', when the fluid is flowing due to density gradients (buoyancy force), or ''forced'', when the fluid is flowing because an external process force it to flow or make the control volume move through the fluid. When air is the considered fluid, convection is always associated to '''conduction''' heat transfer. The reason is that a film is formed on the surface of the object, called the [https://en.wikipedia.org/wiki/Boundary_layer boundary layer], and it is partly steady and at a temperature closer than the temperature of the object than the temperature of the fluid. In that case, conduction applies. An insulation layer's effect can be easily calculated as its thermal [https://en.wikipedia.org/wiki/U-value#U-value transmittance] (U-value) or [https://en.wikipedia.org/wiki/R-value_(insulation) resistance] (R-value), the amount of heat that it allows to be transferred through it. This is directly obtained from the material's [https://en.wikipedia.org/wiki/Thermal_conductivity thermal conductivity] ''k'' (unit: W/m.K) and the thickness of the insulation ''L''. ''R = L/k'' and ''U = k/L''. Unit of U is W/m^2.K. The transmitter heat is then Φ = A × U × (T1 - T2), in Watt (= Joule/s), where A is the external area of the insulation layer, T1 and T2 are the internal and external temperature. Examples for the transmittance of insulation layers can be found here [http://bmeweb.niu.edu.tw/pcwu/%E7%BF%92%E9%A1%8C%E8%A7%A3%E7%AD%94/Heat%20Chap01-087.doc], ===Heat transfer coefficient ''h''=== The rate of heat loss of a body by convection is proportional to the difference in temperatures between the body and its surroundings, as stated by [https://en.wikipedia.org/wiki/Convective_heat_transfer#Newton.27s_law_of_cooling Newton's law of cooling]: {{SERVER}}/images/formulas_mirror/newtons_law_of_cooling_neg.png , where ''h'' is the [https://en.wikipedia.org/wiki/Heat_transfer_coefficient heat transfer coefficient]. This ''h'' depends on many parameters (flow rate, surface roughness, fluid properties, and others) and is very hard to calculate accurately. Approximations exist for some conditions and determining them is still an active research topic for some conditions. It will be our main problem in the case of air to aircraft heat transfer. ===Approximations and conditions=== ''h'' should be calculated from Computational fluid dynamics analysis, which requires a good expertise in the domain and complex software to be used. Since we don't have this capability, we explore existing [https://en.wikipedia.org/wiki/Heat_transfer_coefficient approximations] of the value. These approximations are only valid in some very specific conditions, each condition having a different approximation. Assumptions are also made, for example the temperature of the fluid and the body over the area of the heat transfer are assumed constant, as well as the flow rate. Approximation are often based on a fluid temperature equal to the arithmetic mean between the wall and the free stream. This is called the ''mean film temperature''. ===The Nusselt number ''Nu''=== The [https://en.wikipedia.org/wiki/Nusselt_number Nusselt number] ''Nu'' is [https://en.wikipedia.org/wiki/Heat_transfer_coefficient introduced] when equating Newton's equation to the conduction heat: {{SERVER}}/images/formulas_mirror/nusselt_neg.png = Nu The Nusselt number is then the ratio of the temperature gradient at the surface to the reference temperature gradient, meaning that its value indicates the shape of the temperature gradient. ===The Prandtl number ''Pr''=== The [https://en.wikipedia.org/wiki/Prandtl_number Prandtl number] depends only on the fluid and its state, not on a characteristic length. {{SERVER}}/images/formulas_mirror/prandlt_number_neg.png The Prandtl number controls the relative thickness of the momentum and thermal boundary layers. When Pr is small, it means that the heat diffuses very quickly compared to the velocity (momentum). This means that for liquid metals the thickness of the thermal boundary layer is much bigger than the velocity boundary layer. See page 223 of [2] for other interpretations of the values. ===The Reynolds number ''Re''=== The [https://en.wikipedia.org/wiki/Reynolds_number Reynolds number] gives a measure of the ratio of inertial forces to viscous forces. {{SERVER}}/images/formulas_mirror/reynolds_number_neg.png It is also useful because its value can indicate if the flow in the boundary layer is '''laminar''' (value < 350,000), '''turbulent''' (value > 500,000) or transitional between the two, in which case it depends on other factors such as surface roughness. ===The Rayleigh number ''Ra''=== The [https://en.wikipedia.org/wiki/Rayleigh_Number Rayleigh number] is an indicator for natural convection heat transfer. When it is below the critical value for a fluid, heat transfer is primarily in the form of conduction; when it exceeds the critical value, heat transfer is primarily in the form of convection. {{SERVER}}/images/formulas_mirror/rayleigh_number_neg.png α is the [https://en.wikipedia.org/wiki/Thermal_diffusivity thermal diffusivity] {{SERVER}}/images/formulas_mirror/thermal_diffusivity_neg.png ; β is the [https://en.wikipedia.org/wiki/Coefficient_of_thermal_expansion thermal expansion coefficient], for an isobaric process it can be approximated to β = 1/T. ''x'' is the distance from the leading edge. ==Cases of application== We currently use heat transfer to estimate the rate of vaporization of cryogenic propellant on the ground and during the ascent to rocket ignition altitude. The latter could however be easily used to calculate the drag force of the aircraft fuselage on its tank part. In both case we can assume the fluids are incompressible and steady and that the temperatures are constant over the body and in the free stream fluid for a given altitude. The program using the [[Flight_at_high_altitude#Gas_properties_and_altitude|atmospheric model]] and the flight path model to compute the heat transfer during flight is available freely here: [[File:Heat_transfer_to_rocket_tank.c]]. '''''Warning:''''' ''the values presented below are highly dependent on the dimensions of the considered tank (diameter: 0.27m, length: 1.63m, surface: 1.38261m^2), which was given by the [[Rocket:First_approximations#Minimum_mass_evaluation|first approximation]] of the rocket mass program, the climb rates and flight profile (currently based on [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2]'s rate), and the temperature of the propellant (here LOX at 92K).'' '''''Warning:''''' ''the method and results presented here are based on approximations and assumptions, and it may even have not been done in the proper way. Please validate the method used if you have some knowledge on convective heat transfer.'' ===Natural convection for horizontal cryogenic tank=== On the ground, the cryogenic propellant tank undergoes a large temperature difference and since it doesn't move it's the natural convection that is at work, ''if the wind is neglected''. An approximation exists for the Nusselt number in this condition: [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#External_flow.2C_Horizontal_cylinder horizontal cylinder in external flow under natural convection]. {{SERVER}}/images/formulas_mirror/h_natural_conv_external_horiz_cyl_neg.png D is the diameter of the approximated cylinder, in our case we take 0.27m, given by a [[Rocket:First_approximations#Minimum_mass_evaluation|first evaluation]] of the rocket mass program. Air density is derived from the atmospheric model at sea level, corrected to the actual ground temperature. For a 25°C or 298.15 K ground temperature: * the density (rho) is 1.18391 kg/m^3 and at the mean film temperature of -78.07°C or 195.075 K, given the propellant and tank temperature of 92 K: * the viscosity µ is 1.30015e-05 Pa.s (calculated using [http://www.cfd-online.com/Wiki/Sutherland's_law Sutherland's law]) * the kinematic viscosity nu is µ / rho = 1.0982e-5 m^2/s * the specific heat Cp is 1007.68 J/kg.K (approximated using [http://ninova.itu.edu.tr/tr/dersler/ucak-uzay-fakultesi/965/uck-421/ekkaynaklar?g96162 this interpolation]) * the Prandtl number is 0.73862 * the thermal conductivity of air ''k'' is 0.0177375 W/m.K (approximated using [http://physics.tutorvista.com/heat/heat-transfer.html#thermal-conductivity-of-air this interpolation]) * α is thus k/(rho.Cp) = 1.4868e-5 m^2/s * β is approximated to 1/T = 1/195.075 = 5.126e-3 1/K * Ra_D is ((9.80665 * 5.126e-3) / (1.0982e-5 * 1.4868e-5)) * (298.15 - 92) * 0.27^3 = 6.3467e10 * 0.27^3 = 1.24931e+9. * finally, ''h'' is 0.0177375/0.27 * (0.6 + (0.387 * 1.24931e+9^1/6) / (1 + (0.559/0.73862)^9/16)^8/27)^2 '''= 8.20533 W/m^2.K''' We finally have a heat transfer coefficient for air. We can compute the total heat transferred from it, using Newton's law equation [[Heat_transfer#Heat_transfer_coefficient_h|(top)]]. We need to specify the area on which the heat transfer will apply, the approximated cylinder, and the number of seconds during which the heat is transferred. Tanks are a cylinder with two hemispherical end-caps. Since the end-caps still conduct heat, we will include their area and assume it is part of the approximed cylinder's area. Assumed area is then 1.3826m^2. Let's take 600 seconds (10 minutes) for the time spent on the ground between tanks filling and aircraft lift-off. We assume the temperature to be constant on the ground and in the tank while heat is transferred. This is correct since the heat energy feeds a phase change in the propellant (vaporization) and not a temperature increase of some sort. ''dQ/dt'' = ''h'' * ''A'' * (''T_ground'' - ''T_prop'') = 8.20533 * 1.3826 * (298.15 - 92) = 2338.73 J/s. We can now [[Rocket_Main_Tank#Calculating_evaporation_rate|calculate]] the evaporation rate of the propellant with and without insulation. The heat of vaporization for [[LOX]] is 213 kJ/kg. Without insulation, the vaporization rate of LOX is 11g/s, making 6.59kg of LOX evaporated in 600s. With a 10mm [[Rocket_Main_Tank#Thermal_insulation_materials|Aerogel]] insulation, the thermal conductivity being 15 mW/m.K, its U-value being k/L = 0.015/0.010 = 1.5 W/m^2.K, the vaporization rate is 919.5mg/s. For 10 minutes, it goes down to 0.552kg (instead of 6.59 without insulation). The overall heat transfer coefficient U was calculated using [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#Alternative_Method_.28A_simple_method_for_determining_the_overall_heat_transfer_coefficient.29 this] method. ===Forced convection on aircraft fuselage during flight=== Section 5.8 in [2], ''Heat transfer in turbulent flow over a flat plate'', is the closest known answer to this problem. The tank's cylinder can be considered as a flat plate since there is no pressure change in the local y direction. The Colburn analogy is the approximation that applies in this case. It states that the local Nusselt number, assuming constant wall temperature, is Nu_x = 0.0292 Re_x^4/5 Pr^1/3. This equation is however only valid for Re_c < Re_x < 10⁷ and 0.7 < Pr < 100. From what we calculated with the ISA model, Pr varies between 0.73 and 0.75, and the Reynolds number for the characteristic length L of the tank (Re_L) drops from 4.36714e+07 to 6.68455e+06 with altitude. This value of Re confirms that the boundary layer is turbulent. [2] indicates that the Colburn analogy can still be used for Pr up to 10^8, but with some loss of accuracy. We use the following approximation to calculate the average Nusselt number for forced convection in turbulent flows: Nu_L = 0.0366 (Re_L^0.8 - 14500) Pr^1/3 given in [2]. The calculated average Nusselt number for ascent operation is 42353 at sea level and drops to 1037 at 30km altitude. ''h_L'' can then be calculated too, ''h_L'' = Nu_L * ''k''_air / L. It varies from 460.8 on the ground to 9.6 W/m^2.K at 30km altitude. The simple flight path approximation is used to calculate the average ''h'' value for every second of the flight. We also need to calculate the dQ/dt value for each second, since the external temperature changes as the aircraft ascends. The integrated value of heat Q is finally: Q = 1710.94 kJ. '''For [[LOX]], it translates into more than 80kg of oxidizer evaporated during the ascent, when no insulation is used. With an Aerogel insulation layer of 10mm, the mass of evaporated LOX is evaluated to 0.666kg.''' That corresponds to Q = 141.89 kJ. The overall heat transfer coefficient U was calculated using [https://en.wikipedia.org/wiki/Heat_transfer_coefficient#Alternative_Method_.28A_simple_method_for_determining_the_overall_heat_transfer_coefficient.29 this] method. ==References== 1. '''Principles of Heat Transfer''': SI Edition, Frank Kreith, Raj M. Manglik, Mark S. Bohn, 2010. [http://books.google.com/books?id=GnTrJiu9tKwC&pg=PA420 Google books] 2. '''Textbook Of Heat Transfer''' (4th Edition), S.P. Sukhatme, 2006. [http://books.google.com/books?id=-VgAZm6KWrwC Google books] 3. Lecture on '''Advanced Heat Transfer''', Ibrahim Sezai, 2011. Distributed as open courseware, [http://opencourses.emu.edu.tr/course/view.php?id=23 here] A nice wiki page on heat transfer for rocket engines, regenerative cooling in particular: [http://wiki.fubarlabs.org/fubarwiki/Thermal-Analysis-of-Steady-State-Engines.ashx here]. [[Category:Rocket|Heat transfer]] 578f9a653f32f87cbcceea1786f845fddabbdb76 0 85 510 501 2013-08-08T00:14:29Z Vincent 1 /* Introduction to Computational Fluid Dynamics */ links to other resources wikitext text/x-wiki This page has been created to help find resources on CFD, from a beginner level to achieve what we do in this project (nothing so far). =Introduction to Computational Fluid Dynamics= If you are new to CFD, a must read is this book, as recommended by [http://www.cfd-online.com/Books/show_book.php?book_id=3 cfd-online]: '''Computational Fluid Dynamics: The Basics with Applications''', by John David Anderson, 1995. Yes this book is nearly 20 years old, but is a very nice introduction to the topic, the theory doesn't age. A bad quality pdf is wandering on the Web. My next advice is to read the [http://bookboon.com/en/computational-fluid-dynamics-ebook free book] '''Computational Fluid Dynamics''', by Abdulnaser Sayma, 2009. It brings up to speed on modern techniques like finite element and finite volume methods. "The book provides the reader with the theoretical background of basic CFD methods without going into deep detail into the mathematics or numerical algorithms." Another great source of information is the video lectures on CFD from India's E-learning program ([http://nptel.iitm.ac.in/courses/112105045/ NPTEL]), by Prof. S. Chakraborty. If you want to see what CFD looks like on the theory and programming side, a brief overview of a basic method is presented in: [http://www.gputechconf.com/gtcnew/on-demand-gtc.php?sessionTopic=12&select=+#45 A Practical Introduction to Computational Fluid Dynamics on GPUs] from the GPU Tech Conference 2010 (direct link to [http://us.download.nvidia.com/downloads/GTC_Videos/flvs/2058_GTC2010.mp4 slides+voice mp4] and [http://www.nvidia.com/content/GTC-2010/pdfs/2058_GTC2010.pdf presentation's pdf]). This presentation's main topic is using OpenCL to speed up CFD applications, so OpenCL is briefly presented too if you're interested. [http://www.cfd-online.com/Wiki/Main_Page CFD-Wiki], lots of useful pages and references. cafd4ec5b6624f7f8e4807faaa9de60cac23982c 0 86 514 2013-08-25T22:46:31Z Vincent 1 first link wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen: [http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link] (11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. 7c84495b4d3f7df1d32c18600645bdf1c26b6854 515 514 2013-08-25T23:17:24Z Vincent 1 /* Telemetry */ 2nd link, the TM & RF handbook wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). b7a57483b916d08a234b00c55e9036825640b0f8 516 515 2013-08-27T21:50:57Z Vincent 1 /* Telemetry */ ISM wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. 430b38ddba43224b0287733994e15f41b2b47f73 517 516 2013-08-27T22:09:19Z Vincent 1 /* Telemetry */ conversion table wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] a25d7d39199b9a507d86566db4e9687dce505fa0 518 517 2013-08-29T01:43:13Z Vincent 1 /* Telemetry */ example modules wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] 8972540dd095334253c4c6cd0d7119d676e4141e 519 518 2013-08-31T01:46:16Z Vincent 1 /* Telemetry */ modules links wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package) * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27] * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] c288c38b21409a6cbdad840a3e402ea19a11cf72 520 519 2013-09-01T23:33:10Z Vincent 1 /* Telemetry */ radiosonde information wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package) * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27] * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] 2862095f08e465f80ecec3a7f3bdc78b76dca01b 521 520 2013-09-03T00:58:33Z Vincent 1 /* List of emission modules available on the 869MHz ISM band, 500mW power */ xbee pro wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package) * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27] * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW 722587ae404efef287dbd894a676c1e953e30578 522 521 2013-09-03T11:50:35Z Vincent 1 /* List of emission modules available on the 869MHz ISM band, 500mW power */ TX3H wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package) * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27] * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW 66db48854278019698e519143ae617ac972c8d09 523 522 2013-09-05T00:54:18Z Vincent 1 /* Telemetry */ CS wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package) * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27] * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW e134dc826e6bd469ada2199f922d043e3d0679aa 524 523 2013-09-25T01:35:02Z Vincent 1 /* List of emission modules available on the 869MHz ISM band, 500mW power */ prices for adeunis products wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW 48c98f45632f25957d881df5cb118ec7fa57798b 527 524 2013-09-26T21:00:54Z Vincent 1 reception equipment wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. 0083f18e0564d00197b26db53413751a8a2445f5 528 527 2013-09-28T22:12:41Z Vincent 1 /* Telemetry */ ITU link wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. fb234e68d757d5cbee6c2d453e2873ba1ec5c3a6 529 528 2013-09-29T13:57:46Z Vincent 1 adding the tracking section wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There is no magical solution. What is generally done is that the aircraft provides its position through telemetry, which is then used to refine the pointing of the tracking antenna. If it's lost at some point, a wider beam antenna is used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] to try to get a position information. However since these antennas have a lower gain, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. f8999444f2e19e4e07067a568d4724c7cccdf6ca 530 529 2013-09-29T14:45:39Z Vincent 1 /* Telemetry */ SO50 wikitext text/x-wiki =Telemetry= An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There is no magical solution. What is generally done is that the aircraft provides its position through telemetry, which is then used to refine the pointing of the tracking antenna. If it's lost at some point, a wider beam antenna is used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] to try to get a position information. However since these antennas have a lower gain, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. b6447e2f4e8b1142eb9b2edac6da5b1729a31c8a 531 530 2013-09-29T15:11:44Z Vincent 1 /* Telemetry */ resources section and various fixes wikitext text/x-wiki =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There is no magical solution. What is generally done is that the aircraft provides its position through telemetry, which is then used to refine the pointing of the tracking antenna. If it's lost at some point, a wider beam antenna is used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] to try to get a position information. However since these antennas have a lower gain, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). a8f8bd491c57f5a103a62ed94aa47364ea863dd8 532 531 2013-09-30T01:43:15Z Vincent 1 /* Flying object tracking */ RSSI and another solution for tracking: triangulation wikitext text/x-wiki =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). 608f1c4683922ecc2f278edbd38e85be9ba1696d 536 532 2013-10-01T17:28:07Z Vincent 1 /* Telemetry */ project horus 25mW wikitext text/x-wiki =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). e5f504de88be157c2379e68994a872c9e54c3b93 537 536 2013-10-02T01:16:11Z Vincent 1 /* Reception equipment for the 869MHz band */ antennas section wikitext text/x-wiki =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). 5a56c7704a873dd121f47574c8723e6201ee104f 538 537 2013-10-05T21:41:42Z Vincent 1 /* Telemetry */ 172km RC kit wikitext text/x-wiki =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). A [http://tienda.dmd.es/epages/ea0697.sf/en_GB/?ObjectPath=/Shops/ea0697/Products/PACKLRSULR1/SubProducts/PACKLRSULR1-0001 RC kit] is announced for 172km range with a 500mW module on 869MHz band. Antennas used for this range are a omnidirectional antenna (+5dBi) for the reception and a patch antenna (+9dBi) for the transmission. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). 348de20ce9ba75f0f4883cfd6a4f39f44247ccf4 539 538 2013-10-19T21:42:13Z Vincent 1 /* List of emission modules available on the 869MHz ISM band, 500mW power */ HOPE RF module wikitext text/x-wiki =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). A [http://tienda.dmd.es/epages/ea0697.sf/en_GB/?ObjectPath=/Shops/ea0697/Products/PACKLRSULR1/SubProducts/PACKLRSULR1-0001 RC kit] is announced for 172km range with a 500mW module on 869MHz band. Antennas used for this range are a omnidirectional antenna (+5dBi) for the reception and a patch antenna (+9dBi) for the transmission. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.hoperf.com/rf/module/fsk/RFM12BP.htm HOPE RF RFM12BP] SPI-controlled 500mW module * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). 9b9fe3a350d1997b3676087a805cfc6f2a0d77a2 540 539 2013-11-10T10:55:23Z Vincent 1 /* Reception equipment for the 869MHz band */ dave about reception of RFM22B wikitext text/x-wiki =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper. With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/]: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the telemetry link. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). A [http://tienda.dmd.es/epages/ea0697.sf/en_GB/?ObjectPath=/Shops/ea0697/Products/PACKLRSULR1/SubProducts/PACKLRSULR1-0001 RC kit] is announced for 172km range with a 500mW module on 869MHz band. Antennas used for this range are a omnidirectional antenna (+5dBi) for the reception and a patch antenna (+9dBi) for the transmission. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.hoperf.com/rf/module/fsk/RFM12BP.htm HOPE RF RFM12BP] SPI-controlled 500mW module * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. It is said [http://www.daveakerman.com/?p=277#comments here] that using for reception low cost modules such as those listed above for emission will not allow a long range. Amateur radios should be used instead. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). 345391ecd38b5d58697b8e2ece04c080fe2a51bb 541 540 2013-11-28T01:55:58Z Vincent 1 quick update wikitext text/x-wiki The goal of this page is to give the basics of radio frequency links used for telemetry of amateur high altitude balloons, UAVs, rockets and even low earth orbit satellites, It starts from scratch and should be understandable by anyone needing RF telemetry, and some low-cost open source solutions or designs will be presented for easy reuse. =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain (the gain is function of directivity) are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper, or more massively available, thanks to Wi-Fi for example (2.4GHz band). With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/], which vary for each country, so check your local restrictions. Here is a quick overview: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the link. Modules can be found under $100, depending on what level of capabilities you need. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). A more complete list has been made [[Telemetry#List of emission modules available on the 869MHz ISM band, 500mW power|below]] for the 869.5MHz licence-free band. Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. High altitute balloons (HAB) are also launched a few times a month by amateurs for fun, to get pictures or data from the high atmosphere. In UK, they are [http://www.daveakerman.com/?p=592 most often] transmitting on the 434MHz band, with a [http://www.radiometrix.com/content/ntx2 10mW module]. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). A [http://tienda.dmd.es/epages/ea0697.sf/en_GB/?ObjectPath=/Shops/ea0697/Products/PACKLRSULR1/SubProducts/PACKLRSULR1-0001 RC kit] is announced for 172km range with a 500mW module on 869MHz band. Antennas used for this range are a omnidirectional antenna (+5dBi) for the reception and a patch antenna (+9dBi) for the transmission. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers - ''to be verified''. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Tens of such amateur radio satellites have been launched in the end of November and beginning of December 2013, there are number of accessible opportunities to try to receive messages from space! Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.hoperf.com/rf/module/fsk/RFM12BP.htm HOPE RF RFM12BP] SPI-controlled 500mW module * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html Radiometrix TX3H transmitter], requires coding circuitry, 450mW ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. It is said [http://www.daveakerman.com/?p=277#comments here] that using for reception low cost modules such as those listed above for emission will not allow a long range. Amateur radios should be used instead. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). ===HAB (high altitude balloons) links=== ''All links below to be commented.'' * http://projecthab.co.uk/ * http://chris-stubbs.co.uk/wp/ * http://www.daveakerman.com/ * http://maxdarham.com/Telemetry-Weather-Balloon * http://nerdsville.blogspot.co.uk/ * http://ukhas.org.uk/guides:tracking_guide * http://projecthorus.org/index.php/tracking/ * http://aa1zb.net/Antennas/HighAltitude/HighAltAntennas.html ===SDR (software-defined radio) links=== * http://www.funcubedongle.com/ * https://sdr.osmocom.org/trac/wiki/rtl-sdr * http://jeffskinnerbox.wordpress.com/2013/05/26/rtl-sdr-software-defined-radio-sdr-for-20/ ===Amateur radio satellites=== * http://www.pe0sat.vgnet.nl * http://funcube.org.uk/ ed67710213f9436b92b3f4f2ce13c5d561752806 544 541 2013-12-29T20:32:11Z Vincent 1 Radiometrix BiM3H wikitext text/x-wiki The goal of this page is to give the basics of radio frequency links used for telemetry of amateur high altitude balloons, UAVs, rockets and even low earth orbit satellites, It starts from scratch and should be understandable by anyone needing RF telemetry, and some low-cost open source solutions or designs will be presented for easy reuse. =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain (the gain is function of directivity) are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper, or more massively available, thanks to Wi-Fi for example (2.4GHz band). With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/], which vary for each country, so check your local restrictions. Here is a quick overview: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the link. Modules can be found under $100, depending on what level of capabilities you need. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). A more complete list has been made [[Telemetry#List of emission modules available on the 869MHz ISM band, 500mW power|below]] for the 869.5MHz licence-free band. Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. High altitute balloons (HAB) are also launched a few times a month by amateurs for fun, to get pictures or data from the high atmosphere. In UK, they are [http://www.daveakerman.com/?p=592 most often] transmitting on the 434MHz band, with a [http://www.radiometrix.com/content/ntx2 10mW module]. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). A [http://tienda.dmd.es/epages/ea0697.sf/en_GB/?ObjectPath=/Shops/ea0697/Products/PACKLRSULR1/SubProducts/PACKLRSULR1-0001 RC kit] is announced for 172km range with a 500mW module on 869MHz band. Antennas used for this range are a omnidirectional antenna (+5dBi) for the reception and a patch antenna (+9dBi) for the transmission. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers - ''to be verified''. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Tens of such amateur radio satellites have been launched in the end of November and beginning of December 2013, there are number of accessible opportunities to try to receive messages from space! Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.hoperf.com/rf/module/fsk/RFM12BP.htm HOPE RF RFM12BP] SPI-controlled 500mW module * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * Radiometrix [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html TX3H] 450mW transmitter, requires coding circuitry, and the similar 400mW transmitter/receiver [http://www.radiometrix.com/content/bim3h BiM3H] module. ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. It is said [http://www.daveakerman.com/?p=277#comments here] that using for reception low cost modules such as those listed above for emission will not allow a long range. Amateur radios should be used instead. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). ===HAB (high altitude balloons) links=== ''All links below to be commented.'' * http://projecthab.co.uk/ * http://chris-stubbs.co.uk/wp/ * http://www.daveakerman.com/ * http://maxdarham.com/Telemetry-Weather-Balloon * http://nerdsville.blogspot.co.uk/ * http://ukhas.org.uk/guides:tracking_guide * http://projecthorus.org/index.php/tracking/ * http://aa1zb.net/Antennas/HighAltitude/HighAltAntennas.html ===SDR (software-defined radio) links=== * http://www.funcubedongle.com/ * https://sdr.osmocom.org/trac/wiki/rtl-sdr * http://jeffskinnerbox.wordpress.com/2013/05/26/rtl-sdr-software-defined-radio-sdr-for-20/ ===Amateur radio satellites=== * http://www.pe0sat.vgnet.nl * http://funcube.org.uk/ e72b8c77837f10f569778703f7294ebe401ff8f9 545 544 2013-12-30T23:44:54Z Vincent 1 hoperf links update wikitext text/x-wiki The goal of this page is to give the basics of radio frequency links used for telemetry of amateur high altitude balloons, UAVs, rockets and even low earth orbit satellites, It starts from scratch and should be understandable by anyone needing RF telemetry, and some low-cost open source solutions or designs will be presented for easy reuse. =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain (the gain is function of directivity) are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper, or more massively available, thanks to Wi-Fi for example (2.4GHz band). With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/], which vary for each country, so check your local restrictions. Here is a quick overview: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the link. Modules can be found under $100, depending on what level of capabilities you need. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). A more complete list has been made [[Telemetry#List of emission modules available on the 869MHz ISM band, 500mW power|below]] for the 869.5MHz licence-free band. Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. High altitute balloons (HAB) are also launched a few times a month by amateurs for fun, to get pictures or data from the high atmosphere. In UK, they are [http://www.daveakerman.com/?p=592 most often] transmitting on the 434MHz band, with a [http://www.radiometrix.com/content/ntx2 10mW module]. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). A [http://tienda.dmd.es/epages/ea0697.sf/en_GB/?ObjectPath=/Shops/ea0697/Products/PACKLRSULR1/SubProducts/PACKLRSULR1-0001 RC kit] is announced for 172km range with a 500mW module on 869MHz band. Antennas used for this range are a omnidirectional antenna (+5dBi) for the reception and a patch antenna (+9dBi) for the transmission. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers - ''to be verified''. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Tens of such amateur radio satellites have been launched in the end of November and beginning of December 2013, there are number of accessible opportunities to try to receive messages from space! Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.hoperf.com/rf/fsk_module/ HOPE RF] [http://www.hoperf.com/rf/fsk_module/RFM12BP.htm RFM12BP] SPI-controlled 500mW module * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * Radiometrix [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html TX3H] 450mW transmitter, requires coding circuitry, and the similar 400mW transmitter/receiver [http://www.radiometrix.com/content/bim3h BiM3H] module. ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. It is said [http://www.daveakerman.com/?p=277#comments here] that using for reception low cost modules such as those listed above for emission will not allow a long range. Amateur radios should be used instead. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). ===HAB (high altitude balloons) links=== ''All links below to be commented.'' * http://projecthab.co.uk/ * http://chris-stubbs.co.uk/wp/ * http://www.daveakerman.com/ * http://maxdarham.com/Telemetry-Weather-Balloon * http://nerdsville.blogspot.co.uk/ * http://ukhas.org.uk/guides:tracking_guide * http://projecthorus.org/index.php/tracking/ * http://aa1zb.net/Antennas/HighAltitude/HighAltAntennas.html ===SDR (software-defined radio) links=== * http://www.funcubedongle.com/ * https://sdr.osmocom.org/trac/wiki/rtl-sdr * http://jeffskinnerbox.wordpress.com/2013/05/26/rtl-sdr-software-defined-radio-sdr-for-20/ ===Amateur radio satellites=== * http://www.pe0sat.vgnet.nl * http://funcube.org.uk/ 395a65acd5526cbf11401b425f86a08f43b6d6fc 546 545 2013-12-31T00:16:36Z Vincent 1 TX3H price wikitext text/x-wiki The goal of this page is to give the basics of radio frequency links used for telemetry of amateur high altitude balloons, UAVs, rockets and even low earth orbit satellites, It starts from scratch and should be understandable by anyone needing RF telemetry, and some low-cost open source solutions or designs will be presented for easy reuse. =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain (the gain is function of directivity) are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper, or more massively available, thanks to Wi-Fi for example (2.4GHz band). With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/], which vary for each country, so check your local restrictions. Here is a quick overview: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the link. Modules can be found under $100, depending on what level of capabilities you need. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). A more complete list has been made [[Telemetry#List of emission modules available on the 869MHz ISM band, 500mW power|below]] for the 869.5MHz licence-free band. Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. High altitute balloons (HAB) are also launched a few times a month by amateurs for fun, to get pictures or data from the high atmosphere. In UK, they are [http://www.daveakerman.com/?p=592 most often] transmitting on the 434MHz band, with a [http://www.radiometrix.com/content/ntx2 10mW module]. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). A [http://tienda.dmd.es/epages/ea0697.sf/en_GB/?ObjectPath=/Shops/ea0697/Products/PACKLRSULR1/SubProducts/PACKLRSULR1-0001 RC kit] is announced for 172km range with a 500mW module on 869MHz band. Antennas used for this range are a omnidirectional antenna (+5dBi) for the reception and a patch antenna (+9dBi) for the transmission. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers - ''to be verified''. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Tens of such amateur radio satellites have been launched in the end of November and beginning of December 2013, there are number of accessible opportunities to try to receive messages from space! Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.hoperf.com/rf/fsk_module/ HOPE RF] [http://www.hoperf.com/rf/fsk_module/RFM12BP.htm RFM12BP] SPI-controlled 500mW module * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * Radiometrix [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html TX3H] 450mW transmitter (30 EUR), requires coding circuitry, and the similar 400mW transmitter/receiver [http://www.radiometrix.com/content/bim3h BiM3H] module. ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. It is said [http://www.daveakerman.com/?p=277#comments here] that using for reception low cost modules such as those listed above for emission will not allow a long range. Amateur radios should be used instead. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). ===HAB (high altitude balloons) links=== ''All links below to be commented.'' * http://projecthab.co.uk/ * http://chris-stubbs.co.uk/wp/ * http://www.daveakerman.com/ * http://maxdarham.com/Telemetry-Weather-Balloon * http://nerdsville.blogspot.co.uk/ * http://ukhas.org.uk/guides:tracking_guide * http://projecthorus.org/index.php/tracking/ * http://aa1zb.net/Antennas/HighAltitude/HighAltAntennas.html ===SDR (software-defined radio) links=== * http://www.funcubedongle.com/ * https://sdr.osmocom.org/trac/wiki/rtl-sdr * http://jeffskinnerbox.wordpress.com/2013/05/26/rtl-sdr-software-defined-radio-sdr-for-20/ ===Amateur radio satellites=== * http://www.pe0sat.vgnet.nl * http://funcube.org.uk/ cf3e82ff9a92e9af743bf5028fd35e08033d455f 548 546 2014-01-21T18:29:26Z Vincent 1 /* SDR (software-defined radio) links */ wikitext text/x-wiki The goal of this page is to give the basics of radio frequency links used for telemetry of amateur high altitude balloons, UAVs, rockets and even low earth orbit satellites, It starts from scratch and should be understandable by anyone needing RF telemetry, and some low-cost open source solutions or designs will be presented for easy reuse. =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain (the gain is function of directivity) are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper, or more massively available, thanks to Wi-Fi for example (2.4GHz band). With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/], which vary for each country, so check your local restrictions. Here is a quick overview: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the link. Modules can be found under $100, depending on what level of capabilities you need. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). A more complete list has been made [[Telemetry#List of emission modules available on the 869MHz ISM band, 500mW power|below]] for the 869.5MHz licence-free band. Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. High altitute balloons (HAB) are also launched a few times a month by amateurs for fun, to get pictures or data from the high atmosphere. In UK, they are [http://www.daveakerman.com/?p=592 most often] transmitting on the 434MHz band, with a [http://www.radiometrix.com/content/ntx2 10mW module]. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). A [http://tienda.dmd.es/epages/ea0697.sf/en_GB/?ObjectPath=/Shops/ea0697/Products/PACKLRSULR1/SubProducts/PACKLRSULR1-0001 RC kit] is announced for 172km range with a 500mW module on 869MHz band. Antennas used for this range are a omnidirectional antenna (+5dBi) for the reception and a patch antenna (+9dBi) for the transmission. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers - ''to be verified''. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Tens of such amateur radio satellites have been launched in the end of November and beginning of December 2013, there are number of accessible opportunities to try to receive messages from space! Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.hoperf.com/rf/fsk_module/ HOPE RF] [http://www.hoperf.com/rf/fsk_module/RFM12BP.htm RFM12BP] SPI-controlled 500mW module * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * Radiometrix [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html TX3H] 450mW transmitter (30 EUR), requires coding circuitry, and the similar 400mW transmitter/receiver [http://www.radiometrix.com/content/bim3h BiM3H] module. ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. It is said [http://www.daveakerman.com/?p=277#comments here] that using for reception low cost modules such as those listed above for emission will not allow a long range. Amateur radios should be used instead. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). ===HAB (high altitude balloons) links=== ''All links below to be commented.'' * http://projecthab.co.uk/ * http://chris-stubbs.co.uk/wp/ * http://www.daveakerman.com/ * http://maxdarham.com/Telemetry-Weather-Balloon * http://nerdsville.blogspot.co.uk/ * http://ukhas.org.uk/guides:tracking_guide * http://projecthorus.org/index.php/tracking/ * http://aa1zb.net/Antennas/HighAltitude/HighAltAntennas.html ===SDR (software-defined radio) links=== * [http://www.funcubedongle.com/ FunCube Dongle Pro+] * RTL-SDR ** https://sdr.osmocom.org/trac/wiki/rtl-sdr ** http://jeffskinnerbox.wordpress.com/2013/05/26/rtl-sdr-software-defined-radio-sdr-for-20/ ** http://spectrum.ieee.org/geek-life/hands-on/a-40-softwaredefined-radio ===Amateur radio satellites=== * http://www.pe0sat.vgnet.nl * http://funcube.org.uk/ 3733709b5c2630cad978135e1e8fa6d80e389434 549 548 2014-01-21T18:41:35Z Vincent 1 /* Reception equipment for the 869MHz band */ FCDPP update wikitext text/x-wiki The goal of this page is to give the basics of radio frequency links used for telemetry of amateur high altitude balloons, UAVs, rockets and even low earth orbit satellites, It starts from scratch and should be understandable by anyone needing RF telemetry, and some low-cost open source solutions or designs will be presented for easy reuse. =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain (the gain is function of directivity) are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper, or more massively available, thanks to Wi-Fi for example (2.4GHz band). With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/], which vary for each country, so check your local restrictions. Here is a quick overview: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the link. Modules can be found under $100, depending on what level of capabilities you need. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). A more complete list has been made [[Telemetry#List of emission modules available on the 869MHz ISM band, 500mW power|below]] for the 869.5MHz licence-free band. Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. High altitute balloons (HAB) are also launched a few times a month by amateurs for fun, to get pictures or data from the high atmosphere. In UK, they are [http://www.daveakerman.com/?p=592 most often] transmitting on the 434MHz band, with a [http://www.radiometrix.com/content/ntx2 10mW module]. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). A [http://tienda.dmd.es/epages/ea0697.sf/en_GB/?ObjectPath=/Shops/ea0697/Products/PACKLRSULR1/SubProducts/PACKLRSULR1-0001 RC kit] is announced for 172km range with a 500mW module on 869MHz band. Antennas used for this range are a omnidirectional antenna (+5dBi) for the reception and a patch antenna (+9dBi) for the transmission. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers - ''to be verified''. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Tens of such amateur radio satellites have been launched in the end of November and beginning of December 2013, there are number of accessible opportunities to try to receive messages from space! Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.hoperf.com/rf/fsk_module/ HOPE RF] [http://www.hoperf.com/rf/fsk_module/RFM12BP.htm RFM12BP] SPI-controlled 500mW module * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * Radiometrix [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html TX3H] 450mW transmitter (30 EUR), requires coding circuitry, and the similar 400mW transmitter/receiver [http://www.radiometrix.com/content/bim3h BiM3H] module. ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. It is said [http://www.daveakerman.com/?p=277#comments here] that using for reception low cost modules such as those listed above for emission will not allow a long range. Amateur radios should be used instead. From ukhas [http://ukhas.org.uk/guides:tracking_guide tracking guide]: "''FUNcube Dongle Pro+ - An alternative SDR receiver. More sensitive than the cheap SDR dongles but more expensive. Sensitivity is similar to the Radio receivers listed above''". The [http://www.funcubedongle.com/?page_id=1073 FCDP+] (FUNcube Dongle Pro+) is receive-only on the 150kHz to 240MHz and 420MHz to 1.9GHz bands, and costs around 170 EUR. The receive bandwidth is however limited to around 170kHz, not allowing spread-spectrum encoding reception. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). ===HAB (high altitude balloons) links=== ''All links below to be commented.'' * http://projecthab.co.uk/ * http://chris-stubbs.co.uk/wp/ * http://www.daveakerman.com/ * http://maxdarham.com/Telemetry-Weather-Balloon * http://nerdsville.blogspot.co.uk/ * http://ukhas.org.uk/guides:tracking_guide * http://projecthorus.org/index.php/tracking/ * http://aa1zb.net/Antennas/HighAltitude/HighAltAntennas.html ===SDR (software-defined radio) links=== * [http://www.funcubedongle.com/ FunCube Dongle Pro+] * RTL-SDR ** https://sdr.osmocom.org/trac/wiki/rtl-sdr ** http://jeffskinnerbox.wordpress.com/2013/05/26/rtl-sdr-software-defined-radio-sdr-for-20/ ** http://spectrum.ieee.org/geek-life/hands-on/a-40-softwaredefined-radio ===Amateur radio satellites=== * http://www.pe0sat.vgnet.nl * http://funcube.org.uk/ 197dac79de2a06dd32da02941ef724d65c84e9e8 550 549 2014-01-21T19:04:24Z Vincent 1 UK HAB links explained wikitext text/x-wiki The goal of this page is to give the basics of radio frequency links used for telemetry of amateur high altitude balloons, UAVs, rockets and even low earth orbit satellites, It starts from scratch and should be understandable by anyone needing RF telemetry, and some low-cost open source solutions or designs will be presented for easy reuse. =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain (the gain is function of directivity) are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper, or more massively available, thanks to Wi-Fi for example (2.4GHz band). With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/], which vary for each country, so check your local restrictions. Here is a quick overview: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the link. Modules can be found under $100, depending on what level of capabilities you need. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). A more complete list has been made [[Telemetry#List of emission modules available on the 869MHz ISM band, 500mW power|below]] for the 869.5MHz licence-free band. Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. High altitute balloons (HAB) are also launched a few times a month by amateurs for fun, to get pictures or data from the high atmosphere. In UK, they are [http://www.daveakerman.com/?p=592 most often] transmitting on the 434MHz band, with a [http://www.radiometrix.com/content/ntx2 10mW module]. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). A [http://tienda.dmd.es/epages/ea0697.sf/en_GB/?ObjectPath=/Shops/ea0697/Products/PACKLRSULR1/SubProducts/PACKLRSULR1-0001 RC kit] is announced for 172km range with a 500mW module on 869MHz band. Antennas used for this range are a omnidirectional antenna (+5dBi) for the reception and a patch antenna (+9dBi) for the transmission. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers - ''to be verified''. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Tens of such amateur radio satellites have been launched in the end of November and beginning of December 2013, there are number of accessible opportunities to try to receive messages from space! Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.hoperf.com/rf/fsk_module/ HOPE RF] [http://www.hoperf.com/rf/fsk_module/RFM12BP.htm RFM12BP] SPI-controlled 500mW module * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * Radiometrix [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html TX3H] 450mW transmitter (30 EUR), requires coding circuitry, and the similar 400mW transmitter/receiver [http://www.radiometrix.com/content/bim3h BiM3H] module. ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. It is said [http://www.daveakerman.com/?p=277#comments here] that using for reception low cost modules such as those listed above for emission will not allow a long range. Amateur radios should be used instead. From ukhas [http://ukhas.org.uk/guides:tracking_guide tracking guide]: "''FUNcube Dongle Pro+ - An alternative SDR receiver. More sensitive than the cheap SDR dongles but more expensive. Sensitivity is similar to the Radio receivers listed above''". The [http://www.funcubedongle.com/?page_id=1073 FCDP+] (FUNcube Dongle Pro+) is receive-only on the 150kHz to 240MHz and 420MHz to 1.9GHz bands, and costs around 170 EUR. The receive bandwidth is however limited to around 170kHz, not allowing spread-spectrum encoding reception. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). ===HAB (high altitude balloons) links=== * '''UK HAB projects''' ** http://projecthab.co.uk/ Steve Smith has developed his own board for telemetry, the [http://projecthab.co.uk/2013/12/18/neu-vayu/ VAYU-NTX] board, based on NTX transmitter and Arduino-style MCU and a [http://ava.upuaut.net/store/index.php?route=product/product&path=59_64&product_id=91 uBlox GPS sensor]. ** http://www.daveakerman.com/ Dave holds multiple altitude records, using previously Arduinos and now Raspberry Pis to communicate on the RTTY 434MHz tracking network that uses the [http://ukhas.org.uk/projects:dl-fldigi dl-fldigi] software. With friends he also has launched a [http://www.daveakerman.com/?p=1469 paper plane] from very high, created a [http://www.daveakerman.com/?p=1412 chase car] computer based on a Raspberry Pi, and many other great things. ** http://chris-stubbs.co.uk/wp/ Chris is also using NTX and RFM22B transmitters to downlink data and images taken from altitude. In particular, he analysed the RFM22B frequency changes against temperature changes [http://chris-stubbs.co.uk/wp/?p=295 here]. * http://aa1zb.net/Antennas/HighAltitude/HighAltAntennas.html This page explains an example balloon RF link budget. ''All links below to be commented.'' * http://maxdarham.com/Telemetry-Weather-Balloon * http://nerdsville.blogspot.co.uk/ * http://ukhas.org.uk/guides:tracking_guide * http://projecthorus.org/index.php/tracking/ ===SDR (software-defined radio) links=== * [http://www.funcubedongle.com/ FunCube Dongle Pro+] * RTL-SDR ** https://sdr.osmocom.org/trac/wiki/rtl-sdr ** http://jeffskinnerbox.wordpress.com/2013/05/26/rtl-sdr-software-defined-radio-sdr-for-20/ ** http://spectrum.ieee.org/geek-life/hands-on/a-40-softwaredefined-radio ===Amateur radio satellites=== * http://www.pe0sat.vgnet.nl * http://funcube.org.uk/ cea9177a4db5b329baffdc29eb0d7d8394c58cb7 551 550 2014-03-29T01:42:51Z Vincent 1 /* List of emission modules available on the 869MHz ISM band, 500mW power */ freakduino wikitext text/x-wiki The goal of this page is to give the basics of radio frequency links used for telemetry of amateur high altitude balloons, UAVs, rockets and even low earth orbit satellites, It starts from scratch and should be understandable by anyone needing RF telemetry, and some low-cost open source solutions or designs will be presented for easy reuse. =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain (the gain is function of directivity) are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper, or more massively available, thanks to Wi-Fi for example (2.4GHz band). With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/], which vary for each country, so check your local restrictions. Here is a quick overview: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the link. Modules can be found under $100, depending on what level of capabilities you need. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). A more complete list has been made [[Telemetry#List of emission modules available on the 869MHz ISM band, 500mW power|below]] for the 869.5MHz licence-free band. Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. High altitute balloons (HAB) are also launched a few times a month by amateurs for fun, to get pictures or data from the high atmosphere. In UK, they are [http://www.daveakerman.com/?p=592 most often] transmitting on the 434MHz band, with a [http://www.radiometrix.com/content/ntx2 10mW module]. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). A [http://tienda.dmd.es/epages/ea0697.sf/en_GB/?ObjectPath=/Shops/ea0697/Products/PACKLRSULR1/SubProducts/PACKLRSULR1-0001 RC kit] is announced for 172km range with a 500mW module on 869MHz band. Antennas used for this range are a omnidirectional antenna (+5dBi) for the reception and a patch antenna (+9dBi) for the transmission. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers - ''to be verified''. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Tens of such amateur radio satellites have been launched in the end of November and beginning of December 2013, there are number of accessible opportunities to try to receive messages from space! Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.hoperf.com/rf/fsk_module/ HOPE RF] [http://www.hoperf.com/rf/fsk_module/RFM12BP.htm RFM12BP] SPI-controlled 500mW module * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * Radiometrix [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html TX3H] 450mW transmitter (30 EUR), requires coding circuitry, and the similar 400mW transmitter/receiver [http://www.radiometrix.com/content/bim3h BiM3H] module. * [http://www.freaklabsstore.com/index.php?main_page=product_info&cPath=22&products_id=211 Freakduino Long Range Wireless] Arduino compatible, a SoC computer with an onboard RF module and low-noise amplifier for the 868/900MHz band, for only $45! ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. It is said [http://www.daveakerman.com/?p=277#comments here] that using for reception low cost modules such as those listed above for emission will not allow a long range. Amateur radios should be used instead. From ukhas [http://ukhas.org.uk/guides:tracking_guide tracking guide]: "''FUNcube Dongle Pro+ - An alternative SDR receiver. More sensitive than the cheap SDR dongles but more expensive. Sensitivity is similar to the Radio receivers listed above''". The [http://www.funcubedongle.com/?page_id=1073 FCDP+] (FUNcube Dongle Pro+) is receive-only on the 150kHz to 240MHz and 420MHz to 1.9GHz bands, and costs around 170 EUR. The receive bandwidth is however limited to around 170kHz, not allowing spread-spectrum encoding reception. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). ===HAB (high altitude balloons) links=== * '''UK HAB projects''' ** http://projecthab.co.uk/ Steve Smith has developed his own board for telemetry, the [http://projecthab.co.uk/2013/12/18/neu-vayu/ VAYU-NTX] board, based on NTX transmitter and Arduino-style MCU and a [http://ava.upuaut.net/store/index.php?route=product/product&path=59_64&product_id=91 uBlox GPS sensor]. ** http://www.daveakerman.com/ Dave holds multiple altitude records, using previously Arduinos and now Raspberry Pis to communicate on the RTTY 434MHz tracking network that uses the [http://ukhas.org.uk/projects:dl-fldigi dl-fldigi] software. With friends he also has launched a [http://www.daveakerman.com/?p=1469 paper plane] from very high, created a [http://www.daveakerman.com/?p=1412 chase car] computer based on a Raspberry Pi, and many other great things. ** http://chris-stubbs.co.uk/wp/ Chris is also using NTX and RFM22B transmitters to downlink data and images taken from altitude. In particular, he analysed the RFM22B frequency changes against temperature changes [http://chris-stubbs.co.uk/wp/?p=295 here]. * http://aa1zb.net/Antennas/HighAltitude/HighAltAntennas.html This page explains an example balloon RF link budget. ''All links below to be commented.'' * http://maxdarham.com/Telemetry-Weather-Balloon * http://nerdsville.blogspot.co.uk/ * http://ukhas.org.uk/guides:tracking_guide * http://projecthorus.org/index.php/tracking/ ===SDR (software-defined radio) links=== * [http://www.funcubedongle.com/ FunCube Dongle Pro+] * RTL-SDR ** https://sdr.osmocom.org/trac/wiki/rtl-sdr ** http://jeffskinnerbox.wordpress.com/2013/05/26/rtl-sdr-software-defined-radio-sdr-for-20/ ** http://spectrum.ieee.org/geek-life/hands-on/a-40-softwaredefined-radio ===Amateur radio satellites=== * http://www.pe0sat.vgnet.nl * http://funcube.org.uk/ 16e85856f6148a0a175c33e104e0288e2f48331f 0 47 534 341 2013-10-01T00:06:51Z Vincent 1 /* Alternate bearings */ Blandon air bearings wikitext text/x-wiki =Bearings and cooling= Rotational speed achieved by the engine will probably be above 40000rpm. At these speeds, regular ball bearings may overheat or suffer from a too fast wear. In real turbine engines, bearings are constantly lubricated by oil jets, which poses problems with regards to oil pressurization and leaks in other parts of the engine. Seals are consequently placed close to bearings to prevent leaks, generally carbon seals. To enforce the seal, oil is contained in a casing with an internal pressure lower than the external pressure built from compressed air. That way, air can enter the oil casing, but oil cannot leak outside, in other parts of the engine. Accessories like oil pumps, pipes, fixations, filters, tanks, heat exchangers and so on, are also required. ==Bearings== ===Ball bearings=== Ball or roller bearings are the obvious way to guide rotating parts. They can handle high mechanical constraints radially or even axially, they are inexpensive and their integration is reasonably simple. [https://en.wikipedia.org/wiki/Silicon_nitride#Bearings Silicon nitride bearings] have lots of improvements over regular metal ball bearings. Balls are more than 60% less heavy, thus having a lower inertia at high speeds, implying a more softer contact with the tracks, allowing longer lifetime or higher reachable speeds. They also require less lubrication. Fortunately, silicon nitride bearings have reached market with a large production, and are not over-expensive. ===Alternate bearings=== Fluid or magnetic bearings should be considered. They allow much higher rotation speeds and lower friction, but have two main drawbacks. At standby state, they release the radial constrain on moving parts. In reduced-size turbomachinery, where rotor and stator have to be adjusted to tens of microns, it seems quite complicated to use those bearings. The second drawback is that they require more external hardware, to pressurize the fluid or to provide magnetic energy. However, magnetic bearings have been demonstrated in [http://books.google.com/books?id=AXtqMugS3TQC&lpg=PP1&pg=PA263#v=onepage&q&f=false this paper] <ref>S. Jana, V. Arun Kumar and M. Ananda. '''5-axes levitation of a rotor towards indigenization of the magnetic bearing technology'''. In ''Air breathing engines and aerospace propulsion: proceedings of NCABE 2004'', november 2004.</ref>, in which axial position accuracy is measured below 150µm for a 4kg rotor at around 2000rpm. The rotor position sensor has a resolution of 2µm per mV. Unfortunately, no indication is given about the resting position of the rotor and how it impacts the clearance between rotor and stator. [https://en.wikipedia.org/wiki/Foil_bearing Foil bearings] are a particular type of fluid bearing, that "Unlike aero or hydrostatic bearings, foil bearings require no external pressurisation system for the working fluid, so the hydrodynamic bearing is self-starting". In [http://b-dig.iie.org.mx/BibDig/P06-0351/pdfs/track-16/GT2006-90791.pdf this other paper] <ref>Hooshang Heshmat, Michael J. Tomaszewski, James F. Walton II. '''Small gas turbine engine operating with high temperature foil bearing'''. In ''proceedings of GT2006 ASME Turbo Expo 2006: Power for land, sea and air'', may 2006.</ref>, a small centrifugal turbojet is built to evaluate the ability of [http://www.miti.cc/products-services.html MiTi]'s product, a foil bearing, to sustain very high rotation speeds (120'000rpm) and high temperature (800°C). The bearing has a low spacing between the rotor's journal and the stator fixation, but it is secured, in this paper, using a ball bearing on the compressor side, where the temperature is low. They planned to make a dual-foil bearing, we'll need to check on that. MiTi also demonstrated a [http://www.miti.cc/newsletters/20_150mm_foil_journal_bearing%20_hybrid_foil_magnetic_bearing.pdf hybrid foil magnetic bearing], that has the advantages of magnetic bearings at low speeds and those of foil bearings at high speeds. Air bearing are used by [http://www.bladonjets.com/products/micro-air-bearings/ Bladon Jets] for example, in their small turbines. ==Use of lubricating oil for cooling== In real-world jet engines, cooling is the primary function of oil when conventional bearings are used, even more important than lubrication. That's well explained in [https://www.youtube.com/watch?v=WAia8PwMvQM this AgentJayz video]. A high flow rate of oil is then required, with a heat exchanger somewhere along the oil path. The other fluid for the heat exchanger can be air from the bypass duct or fuel, but in our highly size-constrained engine environment, we'll probably have to move some of the engine's equipment to the wings. But that will have to be studied after the bearing type has been chosen obviously. ==Oil displacement without external pumping== [https://en.wikipedia.org/wiki/Screw_conveyor Screw pumping] is being assessed as the way to displace the oil throughout the engine. There will probably be two bearings in the engine, both requiring an oil bath and the shaft itself can probably also use a little refreshment. The idea is to use the rotation of the shaft to actually displace the oil without requiring external accessories. That would be a very lightweight solution and perhaps not that hard to implement, since our shaft is not hollow. The principle first has to be verified, then be tested with such high rotation rates in order to verify that the drag generated on the shaft is acceptable. That does not solve the fact that cooling the oil requires external hardware (a heat exchanger), and that sealing is mandatory in the oil inlet and outlet areas, generally where the bearings are. ==Other hardware required for lubrication and bearing cooling== Sensors will be required too, at least for oil temperature and displacement confirmation. Oil temperature informs about the status of the engine's bearings. Oil displacement sensor is required to ensure that there is no problem with the oil/cooling flow in the engine and that the measured temperature is not bogus. Simple oil filters should also be put somewhere on the oil lines to prevent the more obvious failures. ==References== <references /> [[Category:Turbofan|Bearings and cooling]] ebf6d4534b7f9f69d4297387ec5a60be2a05c6b2 0 11 535 338 2013-10-01T00:14:56Z Vincent 1 /* Fixing blades to rotor */ price induction and bladon link wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets or other engines. They are however very difficult to design and manufacture and are thus very expensive. On this page, we will explore how costs can be reduced while still having adequate capabilities for high altitude flight. ==General principles== Lots of information are available on [https://en.wikipedia.org/wiki/Turbofan Wikipedia's page]. General principle is that there is a combustion that puts energy into a gas, this energy is extracted by a turbine, and the turbine drives both the fan that provides thrust and the compression stage that feeds the combustion with oxygen. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake, and so on. The fan provides thrust by creating a massive air flow, and the engine's core also creates thrust by evacuating the high-speed hot combustion gas. In commercial turbofan engines, the fan is generally responsible for 90% of the overall thrust. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [https://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. A BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost, has the fan mounted on it. One stage engines exist and are less complicated and expensive to build, but are much less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan, but this is not an easy task due to this very speed. Commercial engines featuring a gearbox for the turbofan's fan are expected to reach market in 2012. Multi-spooled engines prevent this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan, but are not yet optimal. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas and thus the work that can be extracted by the turbines and overall engine efficiency. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seeing what's inside real engines: [http://www.youtube.com/user/AgentJayZ#p/u/16/giRA01IHexk on youtube]. Thanks AgentJayZ! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down with the cost. ===Shaped core or shaped shaft?=== An obvious but important optimization to reduce cost and complexity of manufacturing is to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to design and produce, with a simple [https://en.wikipedia.org/wiki/Lathe lathe] ([https://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [https://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates, for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a shaft turned in raw metal won't change much on its final mass. Moreover, we may use a turbine-level mechanism embedded in the stator to try to cool it, which would make it hollow. The main mechanical issues are probably how to properly fix the blades on rotor and stator, how to fix the rotor on the stator with little gap, and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compressor blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [https://en.wikipedia.org/wiki/Superalloys superalloys] or are ceramic-coated. It's the inability of blades to withstand heat and work that limit the power of the engine. Indeed, around 70% of the gas provided by the compressor is used only for chamber and turbine cooling, instead of using it to burn more fuel and create more thrust. The compressor and the turbine are not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow driven by the action of rotor blades to form inside the engine, which would decrease the energy of the gas. Stator blades or vanes redirect the airflow on the next stage in the more efficient direction. Highest efficiencies are reached in turbofans when gaps are reduced between rotor blades' tip and the stator, as well as between the stator blades' tip and the rotor. As always, good efficiency means high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. The shape of the blade and the parameters of their cascade also affects the efficiency. A small 5 stage supersonic compressor providing the same pressure rise than a 15 stage subsonic compressor is less efficient, but it may be compensated by the higher thrust-to-weight ratio. Blade geometric design is also very complicated. First turbine engines had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1926, [https://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines Alan A. Griffith] proved that if blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to deserve being built. Airfoils for blade designs allow compressor stages to better increase the static pressure since they create an expander, an increasing area for the air flow to pass through. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C so it will be way off-limits, and even below that, it will deform before melting. And should we mention that blade deformation at high-centrifugal loads, caused by the high rpm, is a very good way to blow the engine off? Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the vanes the turbine, which are, on actual engines, made of a ceramic-coated high-temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge or bleed circulates, enters the blades, and evacuates through small holes in the blades (convective cooling and film cooling). For the rotor blades, the same principle is used, but with compressor air passing inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air or a high speed electric engine for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air and compressor stall, re-ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude. The rotor speed information and altimeter may be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and removed axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we aim. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. See this example of small fan, from [http://www.price-induction.com/site_media/images/dgen-net/technologies/optimisation_de_masse_grand.jpg Price Induction]. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, a simple square-section root is probably enough. Another lead is to create the blade disk and the blades in a single piece. This can be done with modern manufacturing process like electric discharge machining, 5-dof machining or even laser-based 3D printing. Here is an example from Bladon Jets, [http://www.bladonjets.com/technology/blisk/ the BLISK]. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blade-roots are inserted and fixed. Creating a perpendicular blade root is already a challenge. Rotor blades would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting should be considered. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Air bled from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose [[Turbofan:Alternative Designs|alternative turbofan designs]] based on a mixed-flow turbine. Several pages have been created in the [[:Category:Turbofan|Turbofan category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three to five stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while sustaining the combustion. * [[Turbofan:Bearings|Bearings and cooling]]: high speed rotations require adapted bearings and cooling, which may be reused for rotor and even turbine cooling. ===Turbofan design procedure=== # evaluate required thrust (from aircraft mass and lift, but also [[Flight_at_high_altitude|flight characteristics]]) # calculate required mass flow rate for the fan (thust is [https://en.wikipedia.org/wiki/Thrust calculated] from MFR and flow speed) # fix bypass ratio and fan diameter and rpm, thus giving core diameter (BPR may be [[Turbofan:Alternative_Designs#Full_transonic_engine_design_in_a_single_spool_with_2.1_BPR|fixed by design]]) # calculate required power to drive the fan alone # evaluate a gross compressor driving power (refined later)<br />&nbsp; # calculate total power that has to be drawn from the turbine (fan + compressor + losses) # calculate mass flow rate for the combustion alone # calculate mass flow rate for cooling chamber and turbine ## calculate mass flow rate for cooling chamber ## evaluate mass flow rate for cooling turbine to add to the latter ## calculate entropy and fluid parameters at combustor discharge (speed, temperature) ## calculate temperature of turbine vanes and blade and check if it is acceptable ## iterate on item 8.1 until temperature is unacceptable # calculate the number of turbine blades and stages required for this power # calculate compressor discharge pressure and pressure ratio # calculate how many compressor stages are required depending on sonic or supersonic blade design and fix design # calculate compressor driving power # iterate on item 6 until total power varies # design blades for all calculated parameters and re-run at item 6, total power may have changed [[Category:Turbofan]] 37236c1ad77c785b78e825fdb9c5cd7370b6710c 0 1 552 542 2014-05-04T00:45:00Z Vincent 1 /* News */ balloon update wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''December 2013:''''' CFD has been stalled for a few months. In parallel, work on [[Telemetry|RF telemetry]] has started. First prototype will be tested in a high-altitude balloon in August 2014, in order to validate the RF module and the ground receiver. '''''March 2013:''''' Since December, learning CFD has been the main activity, and it will probably remain so next months. [[CFD:Introduction|A page]] has been created to give CFD beginners some interesting links. CFD has a steep learning curve, but learning how to use it will have huge benefits for the project in the long term: * evaluate the lift and drag associated with supersonic wings/fins on the aircraft-launched rocket will enable us to refine the rocket mass ([[File:Rocket_mass.c]]) model * hopefully have a theoretical validation of our high-altitude turbofan [[Turbofan:Alternative_Designs|alternative design]] * simulate different wing profiles at high-altitude subsonic conditions for the carrier aircraft * evaluate aircraft and engine capability on lift-off and early flight conditions while they are both tailored for high altitude flight * refine the heat transfer approximations ([[File:Heat_transfer_to_rocket_tank.c]]) for cryogenics tank vaporization, used in the rocket mass model for tank dimensioning, with a better climb profile input too. '''''December 2012:''''' A simple rocket trajectory model has been made in order to evaluate the trajectory of an aircraft- or balloon-released-rocket. It appears that the Delta V taken for granted for gravity drag for these rockets, around 800 m/s, is quite erroneous, or not possible with the expected overall thrust-to-weight ratio. Without taking into account the aerodynamic effects like lift and drag, for a ratio of around 1.5, the Delta V for gravity drag is at least 1300 m/s. To reduce it, the ratio should be higher, like 3.0, in that case it may be possible to have only 800 m/s Delta V for gravity, but the mass of the engine would be much higher. See [[Rocket:First_approximations#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|a solution]] to this issue. The program evaluating the mass of the rocket has been refined (v3), and the forgotten pipe linking the upper tank to the engine adds approximately 150 kg of wet mass too ([[File:Rocket_mass.c]]). Incoming works will study aerodynamic lift and drag of a simple rocket and of a rocket with small supersonic wings to better evaluate the gravity drag for our mass evaluation. '''''November 2012:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 150 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] have been created to illustrate the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. The program ([[File:Rocket_mass.c]]) modelling the mass of rockets is being refined and the cryogenic propellant vaporization has been evaluated by another program ([[File:Heat_transfer_to_rocket_tank.c]]). Both programs are freely available. '''''August 2012:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012:''''' Study is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. '''''February 2012:''''' Study of basic aerodynamics is under way. More man power is expected in April. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. See [http://gravityloss.wordpress.com/2008/04/21/air-breathers-advantage this article] for more information on the differences of rocket and aircraft propulsion efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several N-Prize teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. In [http://www.youtube.com/watch?v=esgc5W_Ufng this video] (SpaceX guys), here captured at SpaceUP, they don't even predict the use of attitude control outside the atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell ([http://www.jpaerospace.com/ JP Aeroospace]) is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) equivalent to the empty weight of the plane and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Small gas turbine engines exist however, even in an axial design, see [http://www.bladonjets.com/technology/gas-turbines/ Bladon Jets] or [http://www.price-induction.com/en/dgen/caracteristiques/ Price Induction] for a commercial turbofan for very small jets (nice picture of the fan [http://www.price-induction.com/site_media/images/dgen-net/technologies/optimisation_de_masse_grand.jpg here]). Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]], Category page: [[:Category:Rocket|list of pages]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost, but also regarding their efficiency, the Isp they can produce. This is [[Rocket:First_approximations#Effects_of_parameter_changes|especially important]] in an SSTO design. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Turbopumps are used on full-scale commercial rockets, but are very complex to build and design. Xcor has created and demonstrated since 2003 a [http://www.xcor.com/products/pumps/ piston pump] for LOX, which is now used on a 1,500 lb-thrust (6.6 kN) LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory is closely tied to the [[Flight_at_high_altitude#Approaches_overview|initial release parameters]], the [[Rocket:First_approximations|flight parameters]] such as thrust and aerodynamics, and [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control], which depends on sensors and actuators: '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. b8ebed685d30b811e36b1b10d0727dc2b33fbdb3 574 552 2015-12-27T16:17:15Z Vincent 1 /* News */ 2015 update wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''December 2015:''''' CFD is has resumed for a few months. Basic [[Telemetry|RF telemetry]] has been validated using a high-altitude balloon, thanks to the resources of [https://ukhas.co.uk/ ukhas], but for a rocket, a much more important data rate is required. That would either require more directional receiving antennas or a more powerful emitter, which requires an amateur radio license. Most pages on this wiki are outdated or unfinished and would not be useful for somebody familiar with aerospace. There is one page that can be still useful, about [[Rocket:First_approximations|rocket mass estimation]]. A simple program was developed and some graphs are shown at the bottom, to easily compute and visualise which parameters cause the mass of an air-to-orbit rocket to change. '''''December 2013:''''' CFD has been stalled for a few months. In parallel, work on [[Telemetry|RF telemetry]] has started. First prototype will be tested in a high-altitude balloon in August 2014, in order to validate the RF module and the ground receiver. '''''March 2013:''''' Since December, learning CFD has been the main activity, and it will probably remain so next months. [[CFD:Introduction|A page]] has been created to give CFD beginners some interesting links. CFD has a steep learning curve, but learning how to use it will have huge benefits for the project in the long term: * evaluate the lift and drag associated with supersonic wings/fins on the aircraft-launched rocket will enable us to refine the rocket mass ([[File:Rocket_mass.c]]) model * hopefully have a theoretical validation of our high-altitude turbofan [[Turbofan:Alternative_Designs|alternative design]] * simulate different wing profiles at high-altitude subsonic conditions for the carrier aircraft * evaluate aircraft and engine capability on lift-off and early flight conditions while they are both tailored for high altitude flight * refine the heat transfer approximations ([[File:Heat_transfer_to_rocket_tank.c]]) for cryogenics tank vaporization, used in the rocket mass model for tank dimensioning, with a better climb profile input too. '''''December 2012:''''' A simple rocket trajectory model has been made in order to evaluate the trajectory of an aircraft- or balloon-released-rocket. It appears that the Delta V taken for granted for gravity drag for these rockets, around 800 m/s, is quite erroneous, or not possible with the expected overall thrust-to-weight ratio. Without taking into account the aerodynamic effects like lift and drag, for a ratio of around 1.5, the Delta V for gravity drag is at least 1300 m/s. To reduce it, the ratio should be higher, like 3.0, in that case it may be possible to have only 800 m/s Delta V for gravity, but the mass of the engine would be much higher. See [[Rocket:First_approximations#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|a solution]] to this issue. The program evaluating the mass of the rocket has been refined (v3), and the forgotten pipe linking the upper tank to the engine adds approximately 150 kg of wet mass too ([[File:Rocket_mass.c]]). Incoming works will study aerodynamic lift and drag of a simple rocket and of a rocket with small supersonic wings to better evaluate the gravity drag for our mass evaluation. '''''November 2012:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 150 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] have been created to illustrate the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. The program ([[File:Rocket_mass.c]]) modelling the mass of rockets is being refined and the cryogenic propellant vaporization has been evaluated by another program ([[File:Heat_transfer_to_rocket_tank.c]]). Both programs are freely available. '''''August 2012:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012:''''' Study of aerodynamics is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. See [http://gravityloss.wordpress.com/2008/04/21/air-breathers-advantage this article] for more information on the differences of rocket and aircraft propulsion efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several N-Prize teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. In [http://www.youtube.com/watch?v=esgc5W_Ufng this video] (SpaceX guys), here captured at SpaceUP, they don't even predict the use of attitude control outside the atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell ([http://www.jpaerospace.com/ JP Aeroospace]) is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) equivalent to the empty weight of the plane and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Small gas turbine engines exist however, even in an axial design, see [http://www.bladonjets.com/technology/gas-turbines/ Bladon Jets] or [http://www.price-induction.com/en/dgen/caracteristiques/ Price Induction] for a commercial turbofan for very small jets (nice picture of the fan [http://www.price-induction.com/site_media/images/dgen-net/technologies/optimisation_de_masse_grand.jpg here]). Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]], Category page: [[:Category:Rocket|list of pages]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost, but also regarding their efficiency, the Isp they can produce. This is [[Rocket:First_approximations#Effects_of_parameter_changes|especially important]] in an SSTO design. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Turbopumps are used on full-scale commercial rockets, but are very complex to build and design. Xcor has created and demonstrated since 2003 a [http://www.xcor.com/products/pumps/ piston pump] for LOX, which is now used on a 1,500 lb-thrust (6.6 kN) LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory is closely tied to the [[Flight_at_high_altitude#Approaches_overview|initial release parameters]], the [[Rocket:First_approximations|flight parameters]] such as thrust and aerodynamics, and [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control], which depends on sensors and actuators: '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. 2b2061576be2bc8cb56fa21de5c62ee672a025db 576 574 2017-11-06T00:55:28Z Vincent 1 /* News */ SpaceUp wikitext text/x-wiki =N-Prize and reflections on low-cost access to space= This Web site aims to gather research in the field of astronautics, rocketry and other technologies that can be used for the N-Prize competition, and more generally, to put microsatellites in orbit at low cost. It is not an official Web site for the [[N-Prize]], the official being here: http://www.n-prize.com/. The goal of this competition is roughly to reproduce the great achievement of the Sputnik in 1957, but for a 20g satellite and with spending less than £1000. However, this Web site and its associated research will not stop after the contest is over, this is more a long term (should I say lifetime?) project. It is hosted by the Open Technology And Science Knowledge Initiative ([http://otaski.org OTASKI]). I'm not part of a team for the N-Prize, nor did I register one, because I don't really have the expertise and resources to actually build something in time before the deadline of the contest in september 2013. Anyway, if you find this project interesting, [[Join|join]] and [[Guidelines|participate]]! Maybe if we are enough to work on the project, it is possible to make it in time. It is also possible to provide a part of the challenge and join together with another team providing the other part. Other teams have for example been developing satellites, rocket engines, and so on. ==What is the LCAS project?== LCAS, standing for low-cost access to space, aims to provide an '''open and innovative low-cost orbital launch system for very small satellites''' (less than 10 kg). Research has led us to consider using an aircraft for rocket launches (air-to-orbit), the body of the plane being the rocket itself. The rocket, or a part of it, as in any other orbital launch system, would achieve orbit and thus could embed a minimum of science, making optional the use of a real satellite as payload. Since the main constraint is to have low costs, we'll have to design and build the carrier plane first, including its turbofan engines, which is probably the hardest part of the whole project, and as far as we know has never been done by amateurs. We thus currently focus on the turbofan [[Build_a_cheap_turbofan|research and design]], on which depends everything else. We may then consider helping other N-Prize teams if this is done in time, or other similar projects outside the contest, by providing them those engines and help with aircraft design and rocket integration. Some other parts of the aircraft/rocket are also being studied, for example the [[EmbeddedRocketComputer|software control]] and the low-cost [[EmbeddedRocketComputer#Sensors|sensors]] that can be used to render the aircraft autonomous at first, then make the rocket go into space and reach a controlled orbit. ==News== ''News are also available on twitter [https://twitter.com/OTASKI @OTASKI]'' '''''October 2017:''''' A talk was made at SpaceUp about the use of nano-spacecraft as a cheap way to explore the solar system (spoiler: there is no cheap way). See the [[:File:Solar_system_exploration_with_cubesats.pdf|presentation]]. '''''December 2015:''''' CFD is has resumed for a few months. Basic [[Telemetry|RF telemetry]] has been validated using a high-altitude balloon, thanks to the resources of [https://ukhas.co.uk/ ukhas], but for a rocket, a much more important data rate is required. That would either require more directional receiving antennas or a more powerful emitter, which requires an amateur radio license. Most pages on this wiki are outdated or unfinished and would not be useful for somebody familiar with aerospace. There is one page that can be still useful, about [[Rocket:First_approximations|rocket mass estimation]]. A simple program was developed and some graphs are shown at the bottom, to easily compute and visualise which parameters cause the mass of an air-to-orbit rocket to change. '''''December 2013:''''' CFD has been stalled for a few months. In parallel, work on [[Telemetry|RF telemetry]] has started. First prototype will be tested in a high-altitude balloon in August 2014, in order to validate the RF module and the ground receiver. '''''March 2013:''''' Since December, learning CFD has been the main activity, and it will probably remain so next months. [[CFD:Introduction|A page]] has been created to give CFD beginners some interesting links. CFD has a steep learning curve, but learning how to use it will have huge benefits for the project in the long term: * evaluate the lift and drag associated with supersonic wings/fins on the aircraft-launched rocket will enable us to refine the rocket mass ([[File:Rocket_mass.c]]) model * hopefully have a theoretical validation of our high-altitude turbofan [[Turbofan:Alternative_Designs|alternative design]] * simulate different wing profiles at high-altitude subsonic conditions for the carrier aircraft * evaluate aircraft and engine capability on lift-off and early flight conditions while they are both tailored for high altitude flight * refine the heat transfer approximations ([[File:Heat_transfer_to_rocket_tank.c]]) for cryogenics tank vaporization, used in the rocket mass model for tank dimensioning, with a better climb profile input too. '''''December 2012:''''' A simple rocket trajectory model has been made in order to evaluate the trajectory of an aircraft- or balloon-released-rocket. It appears that the Delta V taken for granted for gravity drag for these rockets, around 800 m/s, is quite erroneous, or not possible with the expected overall thrust-to-weight ratio. Without taking into account the aerodynamic effects like lift and drag, for a ratio of around 1.5, the Delta V for gravity drag is at least 1300 m/s. To reduce it, the ratio should be higher, like 3.0, in that case it may be possible to have only 800 m/s Delta V for gravity, but the mass of the engine would be much higher. See [[Rocket:First_approximations#The_gravity_drag_and_overall_thrust-to-weight_ratio_issue|a solution]] to this issue. The program evaluating the mass of the rocket has been refined (v3), and the forgotten pipe linking the upper tank to the engine adds approximately 150 kg of wet mass too ([[File:Rocket_mass.c]]). Incoming works will study aerodynamic lift and drag of a simple rocket and of a rocket with small supersonic wings to better evaluate the gravity drag for our mass evaluation. '''''November 2012:''''' Rocket mass [[Rocket:First_approximations#Our_first_approximation|has been estimated]] to 150 kg. Turbofan engines parameters can now be calculated for a real application: aircraft carrier for air-to-orbit single stage rocket, tailored for pico and [https://en.wikipedia.org/wiki/Miniaturized_satellite#Nanosatellite nanosatellites]. Also, [[Rocket:First_approximations#Effects_of_parameter_changes|graphs]] have been created to illustrate the importance of various rocket design parameters, such as thrust-to-weight ratios, mass ratio, payload mass. The program ([[File:Rocket_mass.c]]) modelling the mass of rockets is being refined and the cryogenic propellant vaporization has been evaluated by another program ([[File:Heat_transfer_to_rocket_tank.c]]). Both programs are freely available. '''''August 2012:''''' A first step in the project realization will be a turbofan's compressor blade manufacturing, in order to validate the manufacturing process suitability and low cost for the turbofan. The first compressor stage prototype has to be designed in this optics. However, that requires having a [[Rocket:First_approximations|first approximation]] of the rocket mass in order to also have an estimation of the aircraft size and mass, from which we can estimate turbofan engine's properties: inlet speed, required thrust, blade length, RPM and so on. Blade manufacturing will mostly rely on a thermocaster that we'll have to design too. '''''May 21, 2012:''''' Boeing [http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_05_21_2012_p25-458597.xml has also announced] its low cost orbital launch system, based on the WhiteKnightTwo carrier craft and a hypersonic air-breathing first and second stages. '''''May 2012:''''' Study of aerodynamics is still heavily under way in order to validate our [[Turbofan:Alternative_Designs|alternate turbofan mode of operation]]. This is the first thing to validate before the project can enter a real engine design phase of the engine, which will in turn allow the plane to be designed. ==How to escape from Earth?== Rockets have been used for more than 50 years to escape the gravity of earth. They are good for three things: create an important thrust, go fast, and burn a large amount of propellant. Indeed, the efficiency of a propulsion engine is measured with specific impulse (''I_sp''), and for rocket engines, it is quite low. However, their engine is the only engine that provide the sufficient thrust to climb up with large speeds and to tear of Earth's gravity. Besides altitude, speed is the most important factor when trying to put an object into orbit. Without it, satellites would fall back down on Earth, even if you climb up at 200 miles. Once again, rocket engines, with their high thrust power can achieve sufficient speed (> 8 km/s) before falling back on Earth. Rocket trajectories generally tend to form a curve nearing the square angle, with the beginning of the flight being orthogonal to Earth and the final direction being parallel to Earth's surface. The reason is that since they achieve ultrasonic speeds very quickly, the dynamic air pressure on their body (mainly the fairing), resulting in drag, becomes quite important. It is more efficient to first escape the low atmosphere, with its 85% of its whole mass below 11km altitude, and then pitch to gain the horizontal speed needed for orbital injection without being slowed down by atmospheric friction. [[Image:Rocket_trajectory.png|center|Rocket trajectory: initial vector is vertical, final is tangent]] That particular point of the cost of escaping the atmosphere made me thought about using an aircraft to launch a rocket from the upper atmosphere, reducing considerably the air pressure, the drag, and improving trajectory and efficiency. Moreover, the specific impulse of a turbofan is around ten times greater than the Isp of a rocket engine, since it uses oxygen from the atmosphere to burn its fuel, and not some on-board oxidizer. See [http://gravityloss.wordpress.com/2008/04/21/air-breathers-advantage this article] for more information on the differences of rocket and aircraft propulsion efficiency. For the N-Prize, the cost of the aircraft could be deducted from the overall price since if it can be reused. I started searching and I found out that Orbital already has developped an [https://en.wikipedia.org/wiki/Air_launch_to_orbit air-to-orbit] launch vehicle, called the [https://www.orbital.com/SpaceLaunch/Pegasus/ Pegasus]. It is able to push onto Low Earth Orbit a payload up to 1,000 lbs (450 kg), and it is launched from a full-sized airplane. My goal is thus to study the feasibility of something similar, at very low price, even for the aircraft. A rocket would still be used for air-to-orbit link because nothing else is able to achieve a speed around 9 km/s before falling back on Earth. Some specific technologies can be used to improve efficiency, as explained below in the [[#The rocket|rocket]] section. Several N-Prize teams are working on using Helium or Hydrogen balloons ([https://en.wikipedia.org/wiki/Rockoon rockoons]) to get to the high atmosphere, up to 35 or 40 km and then launch a rocket. It is a nice solution too, and maybe less expensive in the overall, but balloons are not reusable, suffer from imprecise trajectory due to winds, and provide no initial speed. The initial speed of an aircraft carrier would be quite low too in our first designs, but the potential for a supersonic velocity release is not shut. Single stage to orbit (SSTO) are also a promising research field for low-cost orbiting. In [http://www.youtube.com/watch?v=esgc5W_Ufng this video] (SpaceX guys), here captured at SpaceUP, they don't even predict the use of attitude control outside the atmosphere to avoid expensive guidance actuators. The main idea of SSTO is that the launch system (rocket) ''is'' the payload. It does not even aim to insert a smaller satellite into orbit. ==The aircraft== Some aircraft have been exploring the high atmosphere, around 30km high. Contrary to what one would assume, high flight speeds are not needed, if the weight is kept low. The [http://www.nasa.gov/centers/dryden/news/ResearchUpdate/Helios/index.html Helios] for example, autonomous solar powered aircraft, flights at this altitude at 20km/h. John Powell ([http://www.jpaerospace.com/ JP Aeroospace]) is also researching on high altitude propellers and plans to make it to space using a high altitude base for payload transfer to a bigger plane. He describes it well in this [http://www.youtube.com/watch?v=G7LQuVfDwFQ video] interview. The [https://en.wikipedia.org/wiki/Lockheed_U-2 U-2] is a manned reconnaissance aircraft flying at 21km altitude, cruising at relatively high speeds (690km/h). Those planes are designed with a very long wingspan, and low weight, similar to gliders. Another kind of design is the fighter jet, for example the [https://en.wikipedia.org/wiki/Mig_25 MiG-25] which also was an altitude (amongst other) record breaker. It had two powerful turbojet engines with afterburner, allowing him to reach a service altitude of 20km and a maximum altitude of more than 37km. It however required a thrust (200kN) equivalent to the empty weight of the plane and large amounts of fuel to climb this high. The same is true for the [https://en.wikipedia.org/wiki/Sr-71 SR-71]. These concerns of how high altitude is reached - mainly through high engine power or high lift at subsonic flight - is discussed on the page dedicated to [[Flight at high altitude|high altitude flight]]. Currently, the subsonic way is being studied, both for rocket and the aircraft, since a higher release velocity means that the rocket can be smaller, and consequently the plane too. Nevertheless, we would benefit from speed of the aircraft, speed that wouldn't be needed by the rocket to reach. It is a low speed compared to orbital speed though. Supersonic launch speed would be nice, but very hard to achieve. Currently, only subsonic speed is considered in the project. The major issue with the aircraft is [[Build a cheap turbofan|how to build a £100 turbofan?]] Small turbofan engines exist, but are made for or by the military, so very expensive, very reliable, and their use is restricted to missiles or UAVs. Small gas turbine engines exist however, even in an axial design, see [http://www.bladonjets.com/technology/gas-turbines/ Bladon Jets] or [http://www.price-induction.com/en/dgen/caracteristiques/ Price Induction] for a commercial turbofan for very small jets (nice picture of the fan [http://www.price-induction.com/site_media/images/dgen-net/technologies/optimisation_de_masse_grand.jpg here]). Links on wing or aircraft design related to speed and altitude: [http://forum.avsim.net/topic/328556-u-2-climb-rate/page__st__50#entry1952184]. ===Staging and recovery=== ''Main page for ground-related aircraft operation and return trip: [[Aircraft Mission]].'' Separation from the aircraft is a big concern. If wings and tail are directly mounted on the rocket body and jettisoned, they would not need some guidance or attitude control electronics, just a basic parachute system. If the wings are able to fly without the rocket, two guidance systems are needed: one for the rocket and one to get back the aircraft in one piece for future launches. Keeping the N-Prize in mind, the aircraft part of the space launch system should be reusable, so that it doesn't count in the £1000 limit. In that case, it has to be recovered in good condition, either using a chute and a GPS tracker, or a more complicated autonomous or remotely operated return-to-runway and landing system. ===Guidance=== ''Main page: [[EmbeddedRocketComputer|embedded computer]].'' A satellite navigation system may be used in the plane for position tracking if allowed by their owning entities. Other sensors should be shared with the rocket's embedded computer, if choices made for staging and recovery allow it. Sun position can be a very good and easy indicator of attitude, as well as earth curve recognition. Video camera is likely to be the main sensor, since it can provide lots of information for very low cost (but for high processing power). ==The rocket== ''Main page: [[RocketEngines|rocket engines]], Category page: [[:Category:Rocket|list of pages]].'' Some concerns are emphasized in this section, some choices are made too. A list of concerns and how they are handled by existing engine designs can be found on the [[RocketEngines|rocket engines]] page. For the first approximations of the capabilities and properties of our rocket and rocket engine, for example the minimum weight required to achieve orbit, see the [[Rocket:First_approximations|first approximations]] page. ===Fuel=== ''Main page: [[Propellants]].'' Propellants represent the most important part of the weight of what we have to launch. It should thus be chosen carefully regarding to its cost, but also regarding their efficiency, the Isp they can produce. This is [[Rocket:First_approximations#Effects_of_parameter_changes|especially important]] in an SSTO design. Alcohol has been used in the early ages of rocketry, in the German V-2 for example. It has the advantages to be cheap, and burns quite well. It is not pure, generally used between 75 an 90 percent of volume ratio with water for the rest. That water is used to lower the temperature combustion and to keep the engine cool enough to stay in one piece. Rocket-grade kerosene (RP-1) has been introduced later to replace alcohol, providing a better volume efficiency. Alcohol seems to be a very good low cost solution. RP-1 is still used nowadays, and is 20% more efficient than alcohol with a liquid oxygen (LOX) oxidizer. The next question is thus: should we use some pure alcohol, alcohol/water blend or alcohol/something else blend? '''E85''', a 85 percent alcohol and 15 percent gasoline fuel recently put on the automotive fuel market, makes a promising rocket fuel. Its efficiency should be slightly better than alcohol, still being very cheap, around £0.5 a liter. Alcohol has good regenerative cooling properties but the non-refined 15% hydrocarbon in it [[RocketEngines#Cooling_for_a_LOX.2FE85_engine|may prevent]] to use it as a coolant. E85 has a different air-fuel ratio than gasoline, requiring less oxygen (or more fuel) to burn, which can be a good thing for us since a cheap LOX tank may be heavy, so the smaller the better. ===Oxidizer=== ''Main page: [[Propellants]].'' Liquid Oxygen ([[LOX]]) is the obvious/best choice as oxidizer and for high Isp. However, it has the big drawback of being a cryogenic fluid, implying cryogenics [[Rocket_Main_Tank|storage]], cautious manipulation, all that making it quite expensive and complicated. See the [http://books.google.fr/books?id=1OC8zeol7uMC cryogenic engineering book]. Alternatives are [https://en.wikipedia.org/wiki/Nitrous_oxide#Rocket_motors Nitrous oxide] and [https://en.wikipedia.org/wiki/Hydrogen_peroxide Hydrogen peroxide]. The latter would be better, since it's more dense, but it seems complicated and expensive to have it manufactured at a high concentration. ===Engine=== ''Main page: [[RocketEngines|rocket engines]]'' The pump is also a major concern, especially for cost and chamber pressure capability. Turbopumps are used on full-scale commercial rockets, but are very complex to build and design. Xcor has created and demonstrated since 2003 a [http://www.xcor.com/products/pumps/ piston pump] for LOX, which is now used on a 1,500 lb-thrust (6.6 kN) LOX/kerosene engine. [https://en.wikipedia.org/wiki/Aerospike_engine Aerospike] engines may be considered, although they are more efficient than bell shaped nozzles at low altitudes and that we want to launch from high altitude. See web page on [http://www.k-makris.gr/RocketTechnology/Nozzle_Design/nozzle_design.htm nozzle design]. ===Trajectory=== The trajectory has to be precise enough to get a launch authorization for a specific orbit, and in a more practical way, to have orbital parameters matching the mission requirements. Trajectory is closely tied to the [[Flight_at_high_altitude#Approaches_overview|initial release parameters]], the [[Rocket:First_approximations|flight parameters]] such as thrust and aerodynamics, and [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft) attitude control], which depends on sensors and actuators: '''[[EmbeddedRocketComputer#Sensors|sensors]]''': cameras can probably be used on the rocket to determine position of the sun and the Earth's horizon. That will have to be validated, but even if it only allows launches at specific times with clear skies, it can be acceptable for a low-cost launch system. Accelerometers, digital gyroscopes and a compass are really cheap nowadays and can be used for attitude monitoring too. They will likely be used in the fast attitude control loop and to refine the attitude calculated by the camera system. '''Actuators''': if sensors are available, actuators are different story. Two ways of changing attitude of a rocket are generally used, as fins have no impact in the vacuum of space: * the vector of exhaust gas of the rocket engine can be controlled. This is generally done in modern engines using a gimbal mount and hydraulic actuators, or more in a more innovative way, using electromagnetic actuators like [https://en.wikipedia.org/wiki/Vega_rocket#P80 Vega's P80]. Another solution is to put fins at the output of the engine nozzle. * control jets (also known as the [https://en.wikipedia.org/wiki/Attitude_control_(spacecraft)#Thrusters RCS]) are used to control the attitude of the rocket. It's often the case for roll control, when the steam generated for turbopumps is not reintroduced in engines but used to control rool, as in SpaceX' Merlin engine. For pitch and yaw, it requires an independent system, generally based on mono-propellant thrusters. Both solutions pose complicated design issues on the rocket's or engine's hardware, but are mandatory. This is one of the big differences between sub-orbital and orbital space flight. The trajectory itself is a balance between vertical speed, minimizing drag of the rocket that wastes its Delta V capability, an horizontal speed, required to reach orbit. ==The satellite== Lots of strategies for the tracking of the satellite have been exposed: flashing light device, radioactive, EM emitting, mirrors... It's not really the issue for now, and others (like [http://www.wikisat.org/?p=632 WikiSat]) have been working on it already. 03a72c07d5585e3a8d2aa3043dc9e744a95901bd 0 86 553 551 2014-06-12T22:55:31Z Vincent 1 /* List of emission modules available on the 869MHz ISM band, 500mW power */ the tRF click module wikitext text/x-wiki The goal of this page is to give the basics of radio frequency links used for telemetry of amateur high altitude balloons, UAVs, rockets and even low earth orbit satellites, It starts from scratch and should be understandable by anyone needing RF telemetry, and some low-cost open source solutions or designs will be presented for easy reuse. =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain (the gain is function of directivity) are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper, or more massively available, thanks to Wi-Fi for example (2.4GHz band). With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/], which vary for each country, so check your local restrictions. Here is a quick overview: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the link. Modules can be found under $100, depending on what level of capabilities you need. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). A more complete list has been made [[Telemetry#List of emission modules available on the 869MHz ISM band, 500mW power|below]] for the 869.5MHz licence-free band. Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. High altitute balloons (HAB) are also launched a few times a month by amateurs for fun, to get pictures or data from the high atmosphere. In UK, they are [http://www.daveakerman.com/?p=592 most often] transmitting on the 434MHz band, with a [http://www.radiometrix.com/content/ntx2 10mW module]. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). A [http://tienda.dmd.es/epages/ea0697.sf/en_GB/?ObjectPath=/Shops/ea0697/Products/PACKLRSULR1/SubProducts/PACKLRSULR1-0001 RC kit] is announced for 172km range with a 500mW module on 869MHz band. Antennas used for this range are a omnidirectional antenna (+5dBi) for the reception and a patch antenna (+9dBi) for the transmission. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers - ''to be verified''. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Tens of such amateur radio satellites have been launched in the end of November and beginning of December 2013, there are number of accessible opportunities to try to receive messages from space! Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.hoperf.com/rf/fsk_module/ HOPE RF] [http://www.hoperf.com/rf/fsk_module/RFM12BP.htm RFM12BP] SPI-controlled 500mW module * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * Radiometrix [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html TX3H] 450mW transmitter (30 EUR), requires coding circuitry, and the similar 400mW transmitter/receiver [http://www.radiometrix.com/content/bim3h BiM3H] module. * [http://www.freaklabsstore.com/index.php?main_page=product_info&cPath=22&products_id=211 Freakduino Long Range Wireless] Arduino compatible, a SoC computer with an onboard RF module and low-noise amplifier for the 868/900MHz band, for only $45! * MikroElektronika [http://www.mikroe.com/click/trf/ tRF click] module, features the Telit LE70-868 - 868 MHz, 500mW, SMA connector. Available [http://www.rlx.sk/en/433mhz-868mhz-915mhz-24ghz/2924-trf-click-mikroe-1535-telit-le70-868-868-mhz-transceiver-module.html here] in europe for 47.64 EUR. ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. It is said [http://www.daveakerman.com/?p=277#comments here] that using for reception low cost modules such as those listed above for emission will not allow a long range. Amateur radios should be used instead. From ukhas [http://ukhas.org.uk/guides:tracking_guide tracking guide]: "''FUNcube Dongle Pro+ - An alternative SDR receiver. More sensitive than the cheap SDR dongles but more expensive. Sensitivity is similar to the Radio receivers listed above''". The [http://www.funcubedongle.com/?page_id=1073 FCDP+] (FUNcube Dongle Pro+) is receive-only on the 150kHz to 240MHz and 420MHz to 1.9GHz bands, and costs around 170 EUR. The receive bandwidth is however limited to around 170kHz, not allowing spread-spectrum encoding reception. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). ===HAB (high altitude balloons) links=== * '''UK HAB projects''' ** http://projecthab.co.uk/ Steve Smith has developed his own board for telemetry, the [http://projecthab.co.uk/2013/12/18/neu-vayu/ VAYU-NTX] board, based on NTX transmitter and Arduino-style MCU and a [http://ava.upuaut.net/store/index.php?route=product/product&path=59_64&product_id=91 uBlox GPS sensor]. ** http://www.daveakerman.com/ Dave holds multiple altitude records, using previously Arduinos and now Raspberry Pis to communicate on the RTTY 434MHz tracking network that uses the [http://ukhas.org.uk/projects:dl-fldigi dl-fldigi] software. With friends he also has launched a [http://www.daveakerman.com/?p=1469 paper plane] from very high, created a [http://www.daveakerman.com/?p=1412 chase car] computer based on a Raspberry Pi, and many other great things. ** http://chris-stubbs.co.uk/wp/ Chris is also using NTX and RFM22B transmitters to downlink data and images taken from altitude. In particular, he analysed the RFM22B frequency changes against temperature changes [http://chris-stubbs.co.uk/wp/?p=295 here]. * http://aa1zb.net/Antennas/HighAltitude/HighAltAntennas.html This page explains an example balloon RF link budget. ''All links below to be commented.'' * http://maxdarham.com/Telemetry-Weather-Balloon * http://nerdsville.blogspot.co.uk/ * http://ukhas.org.uk/guides:tracking_guide * http://projecthorus.org/index.php/tracking/ ===SDR (software-defined radio) links=== * [http://www.funcubedongle.com/ FunCube Dongle Pro+] * RTL-SDR ** https://sdr.osmocom.org/trac/wiki/rtl-sdr ** http://jeffskinnerbox.wordpress.com/2013/05/26/rtl-sdr-software-defined-radio-sdr-for-20/ ** http://spectrum.ieee.org/geek-life/hands-on/a-40-softwaredefined-radio ===Amateur radio satellites=== * http://www.pe0sat.vgnet.nl * http://funcube.org.uk/ 6c7545996b16b57f7ace246164074c2fcb1bff9f 554 553 2014-06-12T23:11:02Z Vincent 1 /* List of emission modules available on the 869MHz ISM band, 500mW power */ the ciseco ARF module wikitext text/x-wiki The goal of this page is to give the basics of radio frequency links used for telemetry of amateur high altitude balloons, UAVs, rockets and even low earth orbit satellites, It starts from scratch and should be understandable by anyone needing RF telemetry, and some low-cost open source solutions or designs will be presented for easy reuse. =Telemetry= A good first read is ''An introduction to RF telemetry systems'', by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). To summarize, with the same type of antenna, the higher the frequency the lower the range. We should prefer a 500MHz band to a 2.4GHz for example. However directive antennas with higher gain (the gain is function of directivity) are more practical in higher frequencies because the wavelength is shorter and antennas are sized to the wavelength. They may also be cheaper, or more massively available, thanks to Wi-Fi for example (2.4GHz band). With a high gain antenna, a higher frequency link can reach the same range as a lower frequency link with a unity gain antenna. There are some license-free radio frequency bands available, the ISM bands (Industrial Scientific and Medical) [http://www.rfsolutions.co.uk/RFSblog/radio-module-design-tips/], which vary for each country, so check your local restrictions. Here is a quick overview: * In the ''EU'': 433MHz – up to 10mW power, 868MHz: multiple channels with power output up to 500mW. 2.4GHz with outputs up to 10mW. * In the ''US'': 433MHz up to 1mW output, 315MHz up to 10mW output, 915MHz up to 500mW (with restrictions on protocol – spread spectrum). The ~900MHz band seems more promising since it allows for relatively high power without requiring a license, and the bandwidth will be more than enough if no video relay is considered on the link. Modules can be found under $100, depending on what level of capabilities you need. [http://www.texim-europe.com/product/ARF7736AA-UN Here] is an example of RF module of 500mW on 869MHz ($100), [http://store.jdrones.com/RDF900_Telemetry_Modem_p/rdf900mdm1.htm here] a telemetry module of 1W on 902-928 MHz ($90). A more complete list has been made [[Telemetry#List of emission modules available on the 869MHz ISM band, 500mW power|below]] for the 869.5MHz licence-free band. Weather balloons are launched very often (more than 850 twice a day around the world) and the most used product is the [https://en.wikipedia.org/wiki/Radiosonde Radiosonde] Vaisala RS92 and variants. Their RF output is 200mW for the 1680MHz version, and at least 40mW for the 403MHz version. High altitute balloons (HAB) are also launched a few times a month by amateurs for fun, to get pictures or data from the high atmosphere. In UK, they are [http://www.daveakerman.com/?p=592 most often] transmitting on the 434MHz band, with a [http://www.radiometrix.com/content/ntx2 10mW module]. The balloons from [http://projecthorus.org/ Project Horus] are communicating through a [http://www.radiometrix.com/content/ntx2 25mW module] on the 435MHz band and they are able to get telemetry from the balloons at several tens of km away (see [http://projecthorus.org/index.php/tracking/ how]). A [http://tienda.dmd.es/epages/ea0697.sf/en_GB/?ObjectPath=/Shops/ea0697/Products/PACKLRSULR1/SubProducts/PACKLRSULR1-0001 RC kit] is announced for 172km range with a 500mW module on 869MHz band. Antennas used for this range are a omnidirectional antenna (+5dBi) for the reception and a patch antenna (+9dBi) for the transmission. Amateur radio operators have reported catching signals from those radiosondes several hundreds of kilometres away, so '''we definitely don't need more than the allowed 500mW ISM RF power'''. That may however require a high quality reception station with high gain antennas and low-noise amplifiers - ''to be verified''. [http://www.copenhagensuborbitals.com/ Copenhagen Suborbitals] has an open source approach to rocketry too, and the [http://www.copenhagensuborbitals.com/sapphire.php Sapphire] Telemetry System is avaiable on [https://github.com/csete/stlm GitHub]. They use two 1 Watt links, in bands above 2GHz. Amateur radio satellites can be easily received from the ground, although their transmit power can be quite low. They use 145 MHz and 435 MHz bands in various uplink/downlink [https://en.wikipedia.org/wiki/OSCAR#OSCAR_satellite_communications configurations]. For example, the [https://en.wikipedia.org/wiki/Saudi-OSCAR_50 Saudi-OSCAR 50] satellite uses a 250 mW UHF transmitter with a 1/4 wave antenna on the 435MHz band, and it [https://www.youtube.com/watch?v=mv4K41Ztax8 can be received], with quite some noise, with a low cost radio and a 2.15dBi gain 1/2 wave [http://www.mfjenterprises.com/Product.php?productid=MFJ-1717S antenna]. Tens of such amateur radio satellites have been launched in the end of November and beginning of December 2013, there are number of accessible opportunities to try to receive messages from space! Amateur satellites have to declare their orbit and frequencies to the International Telecommunication Union (ITU). This can be done for free [http://www.spacenews.com/article/satellite-telecom/37411spectrum-cops-advising-small-satellite-owners-of-obligations now]. ==List of emission modules available on the 869MHz ISM band, 500mW power== * TIMWO HP868, also known as ARF7581AA, [http://www.texim-europe.com/getfile.aspx?id=5655 1-page PDF link] * [http://www.d-d-s.nl/fotos-adeunis-rf/arf35-folder.pdf ARF35] (IP65 package), [http://www.voctronics.nl/prijslijst-adeunis.htm price] 650 EUR. * [http://www.d-d-s.nl/fotos-adeunis-rf/ARF27-folder.pdf ARF27], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 67 EUR TX, 30 EUR RX. * [http://www.eagle.co.za/display_product_3013.htm ARF 29], [http://www.voctronics.nl/prijslijst-adeunis.htm price] 115 EUR. * [http://shop.ciseco.co.uk/arf-high-power-radio-transceiver/ ARF - High power radio transceiver (Ciseco R011)] £29, undetermined power. * [http://friendcom.diytrade.com/sdp/609348/4/pd-3140416/3316844.html FC-RF209] * [http://www.alibaba.com/product-gs/513410711/500mW_ISM_Narrow_band_high_speed.html] * [http://www.alibaba.com/product-gs/603338630/500mW_RF_Module_with_SPI.html] * [http://www.alibaba.com/product-gs/635123681/500mw_rf_module_data_transceiver_YS.html Yishi YS-C30L] * [http://szmellow.en.alibaba.com/product/848824108-218422702/500mW_ISM_band_low_cost_RF_module.html Mellow ml808] * [http://www.hoperf.com/rf/fsk_module/ HOPE RF] [http://www.hoperf.com/rf/fsk_module/RFM12BP.htm RFM12BP] SPI-controlled 500mW module * [http://www.digi.com/products/wireless-wired-embedded-solutions/zigbee-rf-modules/point-multipoint-rfmodules/xbee-pro-868 XBee-PRO 868HP], announced for 80km with RF line of sight, but limited to 315mW * Radiometrix [http://radiometrixstore.com/transmitters/tx3h-wide-band-450mw-transmitter-frequency-869-50mhz.html TX3H] 450mW transmitter (30 EUR), requires coding circuitry, and the similar 400mW transmitter/receiver [http://www.radiometrix.com/content/bim3h BiM3H] module. * [http://www.freaklabsstore.com/index.php?main_page=product_info&cPath=22&products_id=211 Freakduino Long Range Wireless] Arduino compatible, a SoC computer with an onboard RF module and low-noise amplifier for the 868/900MHz band, for only $45! * MikroElektronika [http://www.mikroe.com/click/trf/ tRF click] module, features the Telit LE70-868 - 868 MHz, 500mW, SMA connector. Available [http://www.rlx.sk/en/433mhz-868mhz-915mhz-24ghz/2924-trf-click-mikroe-1535-telit-le70-868-868-mhz-transceiver-module.html here] in europe for 47.64 EUR. ==Reception equipment for the 869MHz band== Three kinds of choices are offered to us for reception: * the reception module matching the emission module, some of them are indeed developed and sold together; the advantage is that we know what is the sensitivity of the receptor and we know that it will operate without issue on the same band, * an amateur radio equipment, * a software defined radio equipment (SDR), like the populars [http://www.funcubedongle.com/?page_id=1073 FunCube Dongle Pro+], [http://www.nuand.com/ bladeRF] and the [http://www.kickstarter.com/projects/mossmann/hackrf-an-open-source-sdr-platform hackRF]. SDR allows a large range of frequencies to be received and kind of encoding to be decoded. All the work and control is done by a computer, contrary to amateur radio equipment that does it in hardware. SDR interfaces are generally USB dongles on which an antenna is plugged. In any case, a high gain directive antenna operating in the 869MHz band will be required to pickup the signal that far away, or even send some data upstream. It will need to be directed towards, which can prove difficult when objects are behind clouds or in a not well known orbit. We may need a low noise amplifier too, depending on the chosen reception equipment. It is said [http://www.daveakerman.com/?p=277#comments here] that using for reception low cost modules such as those listed above for emission will not allow a long range. Amateur radios should be used instead. From ukhas [http://ukhas.org.uk/guides:tracking_guide tracking guide]: "''FUNcube Dongle Pro+ - An alternative SDR receiver. More sensitive than the cheap SDR dongles but more expensive. Sensitivity is similar to the Radio receivers listed above''". The [http://www.funcubedongle.com/?page_id=1073 FCDP+] (FUNcube Dongle Pro+) is receive-only on the 150kHz to 240MHz and 420MHz to 1.9GHz bands, and costs around 170 EUR. The receive bandwidth is however limited to around 170kHz, not allowing spread-spectrum encoding reception. ===Antennas=== A general principle about antennas is that the narrower the beam, the higher the gain. Indeed, it would take 100 times more power to cover the whole sphere around an ideally isotropic antenna than with an antenna covering a 1/100th of this sphere, with the same perceived power at the same distance. Wikipedia's article on [https://en.wikipedia.org/wiki/High-gain_antenna High-gain antennas] is quite instructive. Here is a quote: high-gain antennas must be physically large, since according to the diffraction limit, the narrower the beam desired, the larger the antenna must be (measured in wavelengths). [https://en.wikibooks.org/wiki/Communication_Systems/Antennas This wikibooks article] has a large list of antenna types and their main characteristics. [http://www.teletopix.org/4g-lte/all-about-antennas/ This link] is a good introduction too, with most properties of antennas explained. ==Flying object tracking== It may not be easy to track a flying object with a directional antenna, even inside the atmosphere. If it passes behind clouds for example, you lose the ability to track it visually and it may be complicated to find it again later. In the case of a high altitude balloon with clear sky, that can be done easily if winds don't push it hundreds of miles away. Otherwise, it may get behind mountains and the line of sight can be lost if it's not high enough in altitude. There are two main solutions to this problem. The first is to let the aircraft provide its position through the telemetry link, which is then used to refine the pointing of the tracking antenna. The issue with this solution is that bad weather may make the radio link or GPS lock unstable, and still result in failure of the tracking. It also requires the aircraft to know its position quite well, but IMU coupled with GPS should be reliable enough. For rockets however, that may be more complicated to have an accurate location information with amateur sensors. The second solution is to have several ground stations to triangulate the position and speed of the emitter, using [https://en.wikipedia.org/wiki/Received_signal_strength_indication received signal strength indication] (RSSI) and Doppler shift. This technique is often used for tracking indoors. Using the RSSI as control loop input with only one station can be done, but bad weather affects it too, and since there are 4 possible actions (2 on each axis), it would be a guess-and-try type of tracking, with a number of missed information. If the tracking is lost at some point, a wider beam antenna may be used, like a [https://en.wikipedia.org/wiki/Patch_antenna patch antenna] (really lost in that case) to try to get a position information. Since these antennas have a lower gain than highly directional antennas, they may not be able to catch the data correctly, but they can still provide a cone of plausible localization. ==Resources== [http://www.cpcstech.com/dbm-to-watt-conversion-information.htm dBm to Watt conversion table] An introduction to RF telemetry systems, by Gale Allen ([http://mavdisk.mnsu.edu/alleng/communications/DataRadio/p_telemetry.pdf pdf link], 11 pages). A more complete reading is the ''Telemetry Systems Radio Frequency Handbook'', US military document, 2008 ([http://www.wsmr.army.mil/RCCsite/Documents/120-08%20Telemetry%20Systems%20Radio%20Frequency%20Handbook/120-08%20Telemetry%20(TM)%20Radio%20Frequency(%20RF)%20Handbook.pdf pdf link], 133 pages). ===HAB (high altitude balloons) links=== * '''UK HAB projects''' ** http://projecthab.co.uk/ Steve Smith has developed his own board for telemetry, the [http://projecthab.co.uk/2013/12/18/neu-vayu/ VAYU-NTX] board, based on NTX transmitter and Arduino-style MCU and a [http://ava.upuaut.net/store/index.php?route=product/product&path=59_64&product_id=91 uBlox GPS sensor]. ** http://www.daveakerman.com/ Dave holds multiple altitude records, using previously Arduinos and now Raspberry Pis to communicate on the RTTY 434MHz tracking network that uses the [http://ukhas.org.uk/projects:dl-fldigi dl-fldigi] software. With friends he also has launched a [http://www.daveakerman.com/?p=1469 paper plane] from very high, created a [http://www.daveakerman.com/?p=1412 chase car] computer based on a Raspberry Pi, and many other great things. ** http://chris-stubbs.co.uk/wp/ Chris is also using NTX and RFM22B transmitters to downlink data and images taken from altitude. In particular, he analysed the RFM22B frequency changes against temperature changes [http://chris-stubbs.co.uk/wp/?p=295 here]. * http://aa1zb.net/Antennas/HighAltitude/HighAltAntennas.html This page explains an example balloon RF link budget. ''All links below to be commented.'' * http://maxdarham.com/Telemetry-Weather-Balloon * http://nerdsville.blogspot.co.uk/ * http://ukhas.org.uk/guides:tracking_guide * http://projecthorus.org/index.php/tracking/ ===SDR (software-defined radio) links=== * [http://www.funcubedongle.com/ FunCube Dongle Pro+] * RTL-SDR ** https://sdr.osmocom.org/trac/wiki/rtl-sdr ** http://jeffskinnerbox.wordpress.com/2013/05/26/rtl-sdr-software-defined-radio-sdr-for-20/ ** http://spectrum.ieee.org/geek-life/hands-on/a-40-softwaredefined-radio ===Amateur radio satellites=== * http://www.pe0sat.vgnet.nl * http://funcube.org.uk/ a983b18261219d3f6ae9b4e002b90f6770b67585 0 17 555 480 2014-08-09T20:26:20Z Vincent 1 /* Rocket Engine */ new link for liquid engines wikitext text/x-wiki This page is an introduction to rocket engines, explaining and comparing some of its subsystems. Our rocket will be based on regular rocket technology, as explained on this page. However some choices need to be made in order to gauge the feasibility of the project as a whole, in particular to have ideas of the possible dimensions of the rocket. These choices are presented on the page of the [[Rocket:First approximations|first approximations for the rocket]]. Other information and pages about the rocket and its flight can be found in the [[:Category:Rocket|Rocket category]]. =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. A large list of liquid-propellant engines with pictures of various parts and schematics can be seen [http://www.rocketrelics.com/liquid_propulsion.htm here]. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== There are four known ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, generally using a ring fuel injector at the periphery of the injector plate, and acts both by cooling the chamber walls by contact and by isolating the walls from the combustion * '''Regenerative cooling''' is most widely used in rocket engines, since it is the most efficient way to have the chamber not being destroyed by heat. The general principle is to use the fuel, or sometimes the oxidizer, to cool the chamber walls before injecting those propellants into the chamber. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. * '''Ablative cooling''' is based on materials that provide cooling by being gently destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. * '''Radiative cooling''' uses the natural capacity of materials to radiate (in infrared light for example) when they are hot. Doing this, they lose energy, and thus cool. This is efficient in the void of space, and is used as the nozzle cooling method for the SpaceX's Merlin Vacuum nozzle (with regenerative cooling for the chamber). ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, instead of fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine failure. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotary Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Using LOX for film cooling has also been [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=369#EngineDevelopment demonstrated], by Armadillo Aerospace. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. From SpaceX's experience, it seems that regenerative cooling is sufficient to sustain the combustion temperature hitting the walls. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. [[Category:Rocket|Rocket engines in general]] 908cb8c613c0076a77300aa53888d99572ae7848 556 555 2015-01-11T15:28:16Z Vincent 1 /* Cooling */ updating the cooling section wikitext text/x-wiki This page is an introduction to rocket engines, explaining and comparing some of its subsystems. Our rocket will be based on regular rocket technology, as explained on this page. However some choices need to be made in order to gauge the feasibility of the project as a whole, in particular to have ideas of the possible dimensions of the rocket. These choices are presented on the page of the [[Rocket:First approximations|first approximations for the rocket]]. Other information and pages about the rocket and its flight can be found in the [[:Category:Rocket|Rocket category]]. =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. A large list of liquid-propellant engines with pictures of various parts and schematics can be seen [http://www.rocketrelics.com/liquid_propulsion.htm here]. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== Without extracting or reducing the heat in some way, no material could sustain the heat of a rocket engine combustion. It's basically just like in turbine engines: the hotter it runs, the higher efficiency it reaches. A trade-off must then be made to have the highest temperature versus what the engineering can withstand. This is probably the main issue of rocket engine development. The Copenhagen Suborbitals [http://copsub.com/the-thermodynamic-ice-bag/ page on heat] is very well documented, amongst others. There are four classic ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, where a film of fuel on the chamber wall acts both by cooling it by contact and by isolating it from the heat of the combustion. It is generally created by a ring of fuel injectors at the periphery of the injector plate. Another kind of film can be created by the deposit of some products on the chamber walls. That has been demonstrated by adding 1 percent of silicone oil in ethanol or 10 percent of ethyl silicate in methanol, which create a SiO2 insulation layer constantly ablating and rebuilding on the walls, but also happens with the carbon (soot) deposited by incomplete combustion of Kerosene-like fuels. * '''Regenerative cooling''' is most widely used in rocket engines, since it is enables very long-duration burns without a large loss in efficiency or performance. The general principle is to use the fuel, or [[#Cooling_for_a_LOX.2FE85_engine|sometimes]] the oxidizer, to cool the chamber walls before injecting it into the chamber to be burned. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. Special care must be taken to ensure that the fuel will not evaporate in large proportion before reaching the injector orifice, will not be transformed too much in the cooling lines (coking or autoignition) and that the injector discharge pressure will match a pressure compatible with the chamber pressure at nominal temperature. * '''Ablative cooling''' is based on materials that provide cooling by gently being destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [https://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. The main downside is that such engine cannot be flown several times, but that's generally not the case anyway. * '''Radiative cooling''' uses the natural capacity of materials to radiate (emit light) when they are hot. Doing this, they lose energy, and thus cool. This is most efficient in the cold of space, and is used as the nozzle cooling method for SpaceX [https://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1D Merlin 1D] (which uses regenerative cooling for the chamber). Additionally, '''Adjusting the Oxidizer/Fuel ratio''' with more fuel than in the stoichiometric ratio will produce a cooler burn, with reduced losses in performance or thrust. Data tables can be found for various fuel combinations. For example, LH2/LOX engines tend to burn at a O/F of around 5 instead of 8. In the LH2 case, that also reduces greatly the volume of the fuel tank, because it's very low dense, and the overall mass and size of the vehicle. ===Cooling for a LOX/E85 engine=== For our rocket engine, based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, instead of fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine failure. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotary Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Using LOX for film cooling has also been [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=369#EngineDevelopment demonstrated], by Armadillo Aerospace. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. From SpaceX's experience, it seems that regenerative cooling is sufficient to sustain the combustion temperature hitting the walls. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. [[Category:Rocket|Rocket engines in general]] a455396010a63e78bbb8d0f999597d5b6d60b610 557 556 2015-01-11T15:51:43Z Vincent 1 /* Cooling for a LOX/E85 engine */ agena wikitext text/x-wiki This page is an introduction to rocket engines, explaining and comparing some of its subsystems. Our rocket will be based on regular rocket technology, as explained on this page. However some choices need to be made in order to gauge the feasibility of the project as a whole, in particular to have ideas of the possible dimensions of the rocket. These choices are presented on the page of the [[Rocket:First approximations|first approximations for the rocket]]. Other information and pages about the rocket and its flight can be found in the [[:Category:Rocket|Rocket category]]. =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. A large list of liquid-propellant engines with pictures of various parts and schematics can be seen [http://www.rocketrelics.com/liquid_propulsion.htm here]. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[http://en.wikipedia.org/wiki/SSME SSME] |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== Without extracting or reducing the heat in some way, no material could sustain the heat of a rocket engine combustion. It's basically just like in turbine engines: the hotter it runs, the higher efficiency it reaches. A trade-off must then be made to have the highest temperature versus what the engineering can withstand. This is probably the main issue of rocket engine development. The Copenhagen Suborbitals [http://copsub.com/the-thermodynamic-ice-bag/ page on heat] is very well documented, amongst others. There are four classic ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, where a film of fuel on the chamber wall acts both by cooling it by contact and by isolating it from the heat of the combustion. It is generally created by a ring of fuel injectors at the periphery of the injector plate. Another kind of film can be created by the deposit of some products on the chamber walls. That has been demonstrated by adding 1 percent of silicone oil in ethanol or 10 percent of ethyl silicate in methanol, which create a SiO2 insulation layer constantly ablating and rebuilding on the walls, but also happens with the carbon (soot) deposited by incomplete combustion of Kerosene-like fuels. * '''Regenerative cooling''' is most widely used in rocket engines, since it is enables very long-duration burns without a large loss in efficiency or performance. The general principle is to use the fuel, or [[#Cooling_for_a_LOX.2FE85_engine|sometimes]] the oxidizer, to cool the chamber walls before injecting it into the chamber to be burned. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. Special care must be taken to ensure that the fuel will not evaporate in large proportion before reaching the injector orifice, will not be transformed too much in the cooling lines (coking or autoignition) and that the injector discharge pressure will match a pressure compatible with the chamber pressure at nominal temperature. * '''Ablative cooling''' is based on materials that provide cooling by gently being destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [https://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. The main downside is that such engine cannot be flown several times, but that's generally not the case anyway. * '''Radiative cooling''' uses the natural capacity of materials to radiate (emit light) when they are hot. Doing this, they lose energy, and thus cool. This is most efficient in the cold of space, and is used as the nozzle cooling method for SpaceX [https://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1D Merlin 1D] (which uses regenerative cooling for the chamber). Additionally, '''Adjusting the Oxidizer/Fuel ratio''' with more fuel than in the stoichiometric ratio will produce a cooler burn, with reduced losses in performance or thrust. Data tables can be found for various fuel combinations. For example, LH2/LOX engines tend to burn at a O/F of around 5 instead of 8. In the LH2 case, that also reduces greatly the volume of the fuel tank, because it's very low dense, and the overall mass and size of the vehicle. ===Cooling for a LOX/E85 engine=== For our rocket engine, probably based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, instead of fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine failure. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotary Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Using LOX for film cooling has also been [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=369#EngineDevelopment demonstrated], by Armadillo Aerospace. The [https://en.wikipedia.org/wiki/RM-81_Agena Bell Agena engine] was another example of oxidizer-cooled engine, but using nitric acid (IRFNA, hypergolic) not LOX. It flew hundreds of time. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. From SpaceX's experience, it seems that regenerative cooling is sufficient to sustain the combustion temperature hitting the walls. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. [[Category:Rocket|Rocket engines in general]] 5b22662bdb91ca61faadb1284782373a8ab8f7f5 558 557 2015-01-11T15:54:49Z Vincent 1 SSME https link wikitext text/x-wiki This page is an introduction to rocket engines, explaining and comparing some of its subsystems. Our rocket will be based on regular rocket technology, as explained on this page. However some choices need to be made in order to gauge the feasibility of the project as a whole, in particular to have ideas of the possible dimensions of the rocket. These choices are presented on the page of the [[Rocket:First approximations|first approximations for the rocket]]. Other information and pages about the rocket and its flight can be found in the [[:Category:Rocket|Rocket category]]. =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. A large list of liquid-propellant engines with pictures of various parts and schematics can be seen [http://www.rocketrelics.com/liquid_propulsion.htm here]. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[https://en.wikipedia.org/wiki/SSME SSME] (RS-25) |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |No |No |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |No |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |? |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== Without extracting or reducing the heat in some way, no material could sustain the heat of a rocket engine combustion. It's basically just like in turbine engines: the hotter it runs, the higher efficiency it reaches. A trade-off must then be made to have the highest temperature versus what the engineering can withstand. This is probably the main issue of rocket engine development. The Copenhagen Suborbitals [http://copsub.com/the-thermodynamic-ice-bag/ page on heat] is very well documented, amongst others. There are four classic ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, where a film of fuel on the chamber wall acts both by cooling it by contact and by isolating it from the heat of the combustion. It is generally created by a ring of fuel injectors at the periphery of the injector plate. Another kind of film can be created by the deposit of some products on the chamber walls. That has been demonstrated by adding 1 percent of silicone oil in ethanol or 10 percent of ethyl silicate in methanol, which create a SiO2 insulation layer constantly ablating and rebuilding on the walls, but also happens with the carbon (soot) deposited by incomplete combustion of Kerosene-like fuels. * '''Regenerative cooling''' is most widely used in rocket engines, since it is enables very long-duration burns without a large loss in efficiency or performance. The general principle is to use the fuel, or [[#Cooling_for_a_LOX.2FE85_engine|sometimes]] the oxidizer, to cool the chamber walls before injecting it into the chamber to be burned. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. Special care must be taken to ensure that the fuel will not evaporate in large proportion before reaching the injector orifice, will not be transformed too much in the cooling lines (coking or autoignition) and that the injector discharge pressure will match a pressure compatible with the chamber pressure at nominal temperature. * '''Ablative cooling''' is based on materials that provide cooling by gently being destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [https://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. The main downside is that such engine cannot be flown several times, but that's generally not the case anyway. * '''Radiative cooling''' uses the natural capacity of materials to radiate (emit light) when they are hot. Doing this, they lose energy, and thus cool. This is most efficient in the cold of space, and is used as the nozzle cooling method for SpaceX [https://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1D Merlin 1D] (which uses regenerative cooling for the chamber). Additionally, '''Adjusting the Oxidizer/Fuel ratio''' with more fuel than in the stoichiometric ratio will produce a cooler burn, with reduced losses in performance or thrust. Data tables can be found for various fuel combinations. For example, LH2/LOX engines tend to burn at a O/F of around 5 instead of 8. In the LH2 case, that also reduces greatly the volume of the fuel tank, because it's very low dense, and the overall mass and size of the vehicle. ===Cooling for a LOX/E85 engine=== For our rocket engine, probably based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, instead of fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine failure. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotary Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Using LOX for film cooling has also been [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=369#EngineDevelopment demonstrated], by Armadillo Aerospace. The [https://en.wikipedia.org/wiki/RM-81_Agena Bell Agena engine] was another example of oxidizer-cooled engine, but using nitric acid (IRFNA, hypergolic) not LOX. It flew hundreds of time. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. From SpaceX's experience, it seems that regenerative cooling is sufficient to sustain the combustion temperature hitting the walls. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. [[Category:Rocket|Rocket engines in general]] 91bd33068101f6555af524d7ac880e623ef5a41a 559 558 2015-01-11T16:05:29Z Vincent 1 /* Rocket Engine */ filling holes in table wikitext text/x-wiki This page is an introduction to rocket engines, explaining and comparing some of its subsystems. Our rocket will be based on regular rocket technology, as explained on this page. However some choices need to be made in order to gauge the feasibility of the project as a whole, in particular to have ideas of the possible dimensions of the rocket. These choices are presented on the page of the [[Rocket:First approximations|first approximations for the rocket]]. Other information and pages about the rocket and its flight can be found in the [[:Category:Rocket|Rocket category]]. =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. A large list of liquid-propellant engines with pictures of various parts and schematics can be seen [http://www.rocketrelics.com/liquid_propulsion.htm here]. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[https://en.wikipedia.org/wiki/SSME SSME] (RS-25) |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |None |None |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |None - pressure-fed |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |Film cooling at least |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== Without extracting or reducing the heat in some way, no material could sustain the heat of a rocket engine combustion. It's basically just like in turbine engines: the hotter it runs, the higher efficiency it reaches. A trade-off must then be made to have the highest temperature versus what the engineering can withstand. This is probably the main issue of rocket engine development. The Copenhagen Suborbitals [http://copsub.com/the-thermodynamic-ice-bag/ page on heat] is very well documented, amongst others. There are four classic ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, where a film of fuel on the chamber wall acts both by cooling it by contact and by isolating it from the heat of the combustion. It is generally created by a ring of fuel injectors at the periphery of the injector plate. Another kind of film can be created by the deposit of some products on the chamber walls. That has been demonstrated by adding 1 percent of silicone oil in ethanol or 10 percent of ethyl silicate in methanol, which create a SiO2 insulation layer constantly ablating and rebuilding on the walls, but also happens with the carbon (soot) deposited by incomplete combustion of Kerosene-like fuels. * '''Regenerative cooling''' is most widely used in rocket engines, since it is enables very long-duration burns without a large loss in efficiency or performance. The general principle is to use the fuel, or [[#Cooling_for_a_LOX.2FE85_engine|sometimes]] the oxidizer, to cool the chamber walls before injecting it into the chamber to be burned. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. Special care must be taken to ensure that the fuel will not evaporate in large proportion before reaching the injector orifice, will not be transformed too much in the cooling lines (coking or autoignition) and that the injector discharge pressure will match a pressure compatible with the chamber pressure at nominal temperature. * '''Ablative cooling''' is based on materials that provide cooling by gently being destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [https://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. The main downside is that such engine cannot be flown several times, but that's generally not the case anyway. * '''Radiative cooling''' uses the natural capacity of materials to radiate (emit light) when they are hot. Doing this, they lose energy, and thus cool. This is most efficient in the cold of space, and is used as the nozzle cooling method for SpaceX [https://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1D Merlin 1D] (which uses regenerative cooling for the chamber). Additionally, '''Adjusting the Oxidizer/Fuel ratio''' with more fuel than in the stoichiometric ratio will produce a cooler burn, with reduced losses in performance or thrust. Data tables can be found for various fuel combinations. For example, LH2/LOX engines tend to burn at a O/F of around 5 instead of 8. In the LH2 case, that also reduces greatly the volume of the fuel tank, because it's very low dense, and the overall mass and size of the vehicle. ===Cooling for a LOX/E85 engine=== For our rocket engine, probably based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, instead of fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine failure. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotary Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Using LOX for film cooling has also been [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=369#EngineDevelopment demonstrated], by Armadillo Aerospace. The [https://en.wikipedia.org/wiki/RM-81_Agena Bell Agena engine] was another example of oxidizer-cooled engine, but using nitric acid (IRFNA, hypergolic) not LOX. It flew hundreds of time. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. From SpaceX's experience, it seems that regenerative cooling is sufficient to sustain the combustion temperature hitting the walls. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. [[Category:Rocket|Rocket engines in general]] c2af280026dd93b34700c10c8aa5f218a80f2606 563 559 2015-01-19T23:45:03Z Vincent 1 /* Injectors */ more about pintle injector cooling wikitext text/x-wiki This page is an introduction to rocket engines, explaining and comparing some of its subsystems. Our rocket will be based on regular rocket technology, as explained on this page. However some choices need to be made in order to gauge the feasibility of the project as a whole, in particular to have ideas of the possible dimensions of the rocket. These choices are presented on the page of the [[Rocket:First approximations|first approximations for the rocket]]. Other information and pages about the rocket and its flight can be found in the [[:Category:Rocket|Rocket category]]. =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. A large list of liquid-propellant engines with pictures of various parts and schematics can be seen [http://www.rocketrelics.com/liquid_propulsion.htm here]. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[https://en.wikipedia.org/wiki/SSME SSME] (RS-25) |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |None |None |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |None - pressure-fed |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |Film cooling at least |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== Without extracting or reducing the heat in some way, no material could sustain the heat of a rocket engine combustion. It's basically just like in turbine engines: the hotter it runs, the higher efficiency it reaches. A trade-off must then be made to have the highest temperature versus what the engineering can withstand. This is probably the main issue of rocket engine development. The Copenhagen Suborbitals [http://copsub.com/the-thermodynamic-ice-bag/ page on heat] is very well documented, amongst others. There are four classic ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, where a film of fuel on the chamber wall acts both by cooling it by contact and by isolating it from the heat of the combustion. It is generally created by a ring of fuel injectors at the periphery of the injector plate. Another kind of film can be created by the deposit of some products on the chamber walls. That has been demonstrated by adding 1 percent of silicone oil in ethanol or 10 percent of ethyl silicate in methanol, which create a SiO2 insulation layer constantly ablating and rebuilding on the walls, but also happens with the carbon (soot) deposited by incomplete combustion of Kerosene-like fuels. * '''Regenerative cooling''' is most widely used in rocket engines, since it is enables very long-duration burns without a large loss in efficiency or performance. The general principle is to use the fuel, or [[#Cooling_for_a_LOX.2FE85_engine|sometimes]] the oxidizer, to cool the chamber walls before injecting it into the chamber to be burned. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. Special care must be taken to ensure that the fuel will not evaporate in large proportion before reaching the injector orifice, will not be transformed too much in the cooling lines (coking or autoignition) and that the injector discharge pressure will match a pressure compatible with the chamber pressure at nominal temperature. * '''Ablative cooling''' is based on materials that provide cooling by gently being destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [https://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. The main downside is that such engine cannot be flown several times, but that's generally not the case anyway. * '''Radiative cooling''' uses the natural capacity of materials to radiate (emit light) when they are hot. Doing this, they lose energy, and thus cool. This is most efficient in the cold of space, and is used as the nozzle cooling method for SpaceX [https://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1D Merlin 1D] (which uses regenerative cooling for the chamber). Additionally, '''Adjusting the Oxidizer/Fuel ratio''' with more fuel than in the stoichiometric ratio will produce a cooler burn, with reduced losses in performance or thrust. Data tables can be found for various fuel combinations. For example, LH2/LOX engines tend to burn at a O/F of around 5 instead of 8. In the LH2 case, that also reduces greatly the volume of the fuel tank, because it's very low dense, and the overall mass and size of the vehicle. ===Cooling for a LOX/E85 engine=== For our rocket engine, probably based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, instead of fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine failure. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotary Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Using LOX for film cooling has also been [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=369#EngineDevelopment demonstrated], by Armadillo Aerospace. The [https://en.wikipedia.org/wiki/RM-81_Agena Bell Agena engine] was another example of oxidizer-cooled engine, but using nitric acid (IRFNA, hypergolic) not LOX. It flew hundreds of time. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. The [http://www.theorbitalmechanics.com/show-notes/2015/1/6/episode-3-heat third podcast] of ''[http://www.theorbitalmechanics.com/ The Orbital Mechanics]'' focuses on heat management, and states that pintles are quite good because the cold mixed fuel hits the chamber walls before being ignited, cooling the chamber very efficiently on the upper part. SpaceX success shows that it's quite manageable, though they also use regenerative cooling for the throat and chamber at least. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. [[Category:Rocket|Rocket engines in general]] 75d3211d63e6cbd0d42d84382acd6f694832df07 564 563 2015-01-19T23:47:01Z Vincent 1 /* Cooling */ new link for cooling wikitext text/x-wiki This page is an introduction to rocket engines, explaining and comparing some of its subsystems. Our rocket will be based on regular rocket technology, as explained on this page. However some choices need to be made in order to gauge the feasibility of the project as a whole, in particular to have ideas of the possible dimensions of the rocket. These choices are presented on the page of the [[Rocket:First approximations|first approximations for the rocket]]. Other information and pages about the rocket and its flight can be found in the [[:Category:Rocket|Rocket category]]. =Rocket Engine= The general principle may be simple, but there are numerous ways of achieving it. Different features and properties differ between existing rocket engines, and they all have consequences on complexity of manufacturing, complexity of operation, cost and weight for example. A large list of liquid-propellant engines with pictures of various parts and schematics can be seen [http://www.rocketrelics.com/liquid_propulsion.htm here]. We gather in this table the main properties of some of the existing rocket engines, mostly innovative designs. {| border="1" cellpadding="5" cellspacing="0" |+ align="bottom" |''Rocket engines features'' |- |'''Company''' !Rocketdyne !NPO Energomash !XCOR !XCOR !Armadillo |- |'''Model''' |[https://en.wikipedia.org/wiki/SSME SSME] (RS-25) |[http://www.astronautix.com/engines/rd178d74.htm RD-107] series ([http://www.mentallandscape.com/S_R7.htm Soyuz]) |[http://www.xcor.com/products/engines/4A3_LOX_alcohol_rocket_engine.html XR-4A3 (EZ-rocket)] |[http://www.xcor.com/products/engines/5K18_LOX-kerosene_rocket_engine.html XR-5K18 (Lynx)] |[http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=366 LOX/methane (no name)] |- !colspan="5"|Combustion |- |'''Propellants''' |LOX & LH2 |LOX & Kerosene |LOX & Alcohol |LOX & Kerosene |LOX & LCH4 |- |[[#Pumps and tank pressurization|'''Tank pressurization''']] |Yes, with O2 and H2 gases |Yes, with Nitrogen (same pump than propellants) |None |None |Yes, with Helium |- |[[#Pumps and tank pressurization|'''Fuel pump''']] |Turbopump |Turbopump driven by gaz generator using hydrogen peroxide decomposition (8300rpm) |Piston pump |Piston pump |None - pressure-fed |- |[[#Cooling|'''Cooling''']] |Regenerative w/ LH2 in three stages |Regenerative w/ kerosene (5 mm deep channels milled in the inner wall) and film of kerosene |Regenerative (w/ Alcohol?) |Regenerative w/ Kerosene |Film cooling at least |- |[[#Injectors|'''Injector''']] |? |337 swirling/mixing injectors, ring of kerosene only for film cooling - [[:Image:S_RD107_Head.jpg|view cut]] |? |? |? |- |'''Chamber metal''' |Copper or iron? |6 mm thick chromium bronze alloy inner wall, steel outer wall |Copper |Copper |? |- |'''Ignition system''' |? |Pyrotechnic, soon hypergolic |? |? |? |- !colspan="5"|Actuators |- |'''Energy''' |Hydraulic | | | |Electric |- |'''Provided by''' |Engine's turbopumps | | | |? |- |'''Actuator''' |Six hydraulic servoactuators |Static engine, control by vernier engines |None |None |Servo-motor |- !colspan="5"|Others |- |'''Valves''' |Hydraulically or pneumatically (helium) actuated |? |? |? |? |- |} ==Pumps and tank pressurization== In order to get fuel from the tanks into the combustion chamber, the tanks must be either pressurized or the fuels pumped. In some cases, both techniques are used. The choice for this concern has a large impact on the design of the engine's hardware, and the complexity of manufacturing and operations. Historically, only turbo pumps have been able to feed the engine at a large enough rate (high pressure chamber). Reciprocating pumps have been used in the past, but provided lower pressure and probably more weight. Innovative solutions appeared in research projects or private space projects, like the use of piston pumps for LOX or simple pressurization using liquid helium. Several possibilities exist for tank pressurization: * vaporization of liquid propellants back into their own tanks * external vaporization of inert gas like Helium (can Nitrogen be used for that?) * smoke generator, that basically react fuel and oxidizer and use the resulting smoke for pressurization. The tank design is by itself complicated and now has a specific [[Rocket_Main_Tank|page]]. ==Cooling== Without extracting or reducing the heat in some way, no material could sustain the heat of a rocket engine combustion. It's basically just like in turbine engines: the hotter it runs, the higher efficiency it reaches. A trade-off must then be made to have the highest temperature versus what the engineering can withstand. This is probably the main issue of rocket engine development. The Copenhagen Suborbitals [http://copsub.com/the-thermodynamic-ice-bag/ page on heat] is very well documented, [http://www.theorbitalmechanics.com/show-notes/2015/1/6/episode-3-heat this podcast] from [http://www.theorbitalmechanics.com/ The Orbital Mechanics] is very informative too, amongst others. There are four classic ways to cool a rocket engine: * '''Film cooling''' (''aka'' the cooling curtain) takes place inside the chamber, where a film of fuel on the chamber wall acts both by cooling it by contact and by isolating it from the heat of the combustion. It is generally created by a ring of fuel injectors at the periphery of the injector plate. Another kind of film can be created by the deposit of some products on the chamber walls. That has been demonstrated by adding 1 percent of silicone oil in ethanol or 10 percent of ethyl silicate in methanol, which create a SiO2 insulation layer constantly ablating and rebuilding on the walls, but also happens with the carbon (soot) deposited by incomplete combustion of Kerosene-like fuels. * '''Regenerative cooling''' is most widely used in rocket engines, since it is enables very long-duration burns without a large loss in efficiency or performance. The general principle is to use the fuel, or [[#Cooling_for_a_LOX.2FE85_engine|sometimes]] the oxidizer, to cool the chamber walls before injecting it into the chamber to be burned. The coolant flows into a series of pipes or milling into the external or intermediate walls of the engine, either around the nozzle, the chamber or both of them. Special care must be taken to ensure that the fuel will not evaporate in large proportion before reaching the injector orifice, will not be transformed too much in the cooling lines (coking or autoignition) and that the injector discharge pressure will match a pressure compatible with the chamber pressure at nominal temperature. * '''Ablative cooling''' is based on materials that provide cooling by gently being destroyed, like the heat-shield of spaceships, or the carbon fiber composite nozzle of SpaceX [https://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1A Merlin 1A] engine. The main downside is that such engine cannot be flown several times, but that's generally not the case anyway. * '''Radiative cooling''' uses the natural capacity of materials to radiate (emit light) when they are hot. Doing this, they lose energy, and thus cool. This is most efficient in the cold of space, and is used as the nozzle cooling method for SpaceX [https://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1D Merlin 1D] (which uses regenerative cooling for the chamber). Additionally, '''Adjusting the Oxidizer/Fuel ratio''' with more fuel than in the stoichiometric ratio will produce a cooler burn, with reduced losses in performance or thrust. Data tables can be found for various fuel combinations. For example, LH2/LOX engines tend to burn at a O/F of around 5 instead of 8. In the LH2 case, that also reduces greatly the volume of the fuel tank, because it's very low dense, and the overall mass and size of the vehicle. ===Cooling for a LOX/E85 engine=== For our rocket engine, probably based on LOX and a cheap fuel like E85 or JP-A, we will consider the use of LOX as the coolant, instead of fuel, since cheap fuel polymerizes into cooling pipes, resulting in obstruction and engine failure. LOX as coolant already has been studied by NASA: <blockquote>LOX cooling at chamber pressures to 1500 psia was demonstrated by in-house testing at the NASA Lewis Research Center in the late 1980s. Chambers were fired with cracks to demonstrate wall integrity at elevated LOX mixture ratios. See AIAA paper 89-2739 or NASA TM 10211 3.</blockquote> and by Rotary Rocket and seems feasible as stated [http://yarchive.net/space/launchers/roton.html here] by Doug Jones (Rotary Rocket): <blockquote>"Jet A is a lousy coolant, we have 2.9x the mass of LOX as of fuel available for cooling, and (most important), the LOX has more pressure available for cooling. Bear in mind that flowing through the coolant passages requires a substantial pressure drop, and since the LOX is denser than the fuel, it reaches higher pressure in the centrifugal pumping of the wheel. Thus it is the logical choice for coolant- and it does not foul, no how no way."</blockquote> Using LOX for film cooling has also been [http://www.armadilloaerospace.com/n.x/Armadillo/Home/News?news_id=369#EngineDevelopment demonstrated], by Armadillo Aerospace. The [https://en.wikipedia.org/wiki/RM-81_Agena Bell Agena engine] was another example of oxidizer-cooled engine, but using nitric acid (IRFNA, hypergolic) not LOX. It flew hundreds of time. ==Injectors== Injector role is to mix propellants in the combustion chamber in a way that will produce the most efficient possible combustion. It faces several challenges, such as flow variations, pressure variations in the chamber leading to POGO, film cooling of the chamber walls. It determines the precise start sequence that will not explode the chamber, a process amusingly also called spontaneous disassembly. The temperature of combustion, the combustion ratio, and chamber pressure directly depend on the injector's design. Injectors are most often composed, nowadays and in expensive engines, by hundreds of coaxial fuel/oxidizer injector elements. They assure a combustion efficiency over 99%, so many injector elements mixing very nicely the propellants together. An alternative design comes from the research of TRW in the sixties, and is called the pintle injector design, or pintle engine. In [[:Image:Pintle_engine_paper.pdf|this]] recent paper, TRW summarizes all achievements and the numerous benefits of such engines, which are very interesting for our goal here. Pintle engines only have one injector element, and are thus much less expensive to produce than traditional hundred-elements injectors. They however provide a perfectly stable combustion, with efficiency over 96%, for engines of any scale, with any propellants, and are able to deep throttle up to 1:35. The propellants enter in collision at the exit of the pintle, mixing them efficiently, but requiring more space than in traditional injector design. The Lunar Module Descent Engine is probably the most famous pintle engine, but SpaceX is using them too now. It has never been seen (by TRW at least) that a pintle engine failed or had combustion instability. Bomb tests have always been successful, for any engine size. There may be only three drawbacks to these injector designs: * combustion efficiency is a bit lower than highly complicated injector designs but still good, * combustion chamber requires to be longer than in multiple-injector elements since the combustion is not made parallel to the injector's head but in a torus/cone a bit more distant, * film cooling may be more complicated to design, because there is no specific elements for this purpose. In the [[:Image:Pintle_engine_paper.pdf|paper]], it is said that the film is obtained by pintle tuning. I think it must be difficult to have both a good film and a good combustion efficiency. The [http://www.theorbitalmechanics.com/show-notes/2015/1/6/episode-3-heat third podcast] of ''[http://www.theorbitalmechanics.com/ The Orbital Mechanics]'' focuses on heat management, and states that pintles are quite good because the cold mixed fuel hits the chamber walls before being ignited, cooling the chamber very efficiently on the upper part. SpaceX success shows that it's quite manageable, though they also use regenerative cooling for the throat and chamber at least. ==Pneumatic and hydraulic pressure for actuators and valves== As we can see in the table at the top, different possibilities exist for actuating. The SSME uses hydraulic in nominal mode and pneumatics using He for backup. In satellites, lots of valves are pyrotechnically actuated. Obtaining the pressurization in the system is not easy and is generally done by the fuel pump. SpaceX provided an elegant solution to hydraulic pressure by using the fuel (RP-1) as hydraulic fluid for the launcher, fuel pressurized by the main fuel turbopump. [[Category:Rocket|Rocket engines in general]] 288e2a94392acc644cc0d01136701594b3f9d4bf 0 11 560 535 2015-01-11T21:28:12Z Vincent 1 /* General principles */ updating section wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets or other engines. They are however very difficult to design and manufacture and are thus very expensive. On this page, we will explore how costs can be reduced while still having adequate capabilities for high altitude flight. ==General principles== A lot of information is available on Wikipedia's [https://en.wikipedia.org/wiki/Turbofan turbofan page]. General principle is that there is a combustion that puts energy into a gas, this energy is extracted by a turbine, and the turbine drives both the fan that provides thrust and the compression stage that feeds the combustion with oxygen. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake, and so on. The fan provides thrust by creating a massive air flow, and the engine's core also creates thrust by evacuating the high-speed hot combustion gas. In commercial turbofan engines, the fan is generally responsible for 90% of the overall thrust. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [https://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. An engine with a BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost, has the fan mounted on it. One-stage engines exist and are less complicated and expensive to build, but are much less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan, but this is not an easy task due to its very speed. Commercial engines featuring a gearbox for the turbofan's fan ([https://en.wikipedia.org/wiki/Geared_turbofan geared turbofans]) have reached market in the past few years. Multi-spooled engines lessen this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. High-speed fans are complicated to understand and design, and undergo higher losses due to transonic regime, and having a gearbox on a twin-spool engine is even better, to get a high efficiency turbine and a slowly rotating fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas and thus the work that can be extracted by the turbines. * The turbine inlet temperature is the key parameter for turbine engine. Attaining higher temperatures after combustion, in the turbine inlet, makes an engine more efficient, as explained by the [https://en.wikipedia.org/wiki/Brayton_cycle Brayton thermodynamic cycle]. The obvious issue is the temperature that materials can withstand, in particular the first turbine guide vane and the first high-pressure turbine blades. [https://en.wikipedia.org/wiki/High_temperature_metal High-temperature alloys] and special cooling techniques, like [https://en.wikipedia.org/wiki/Turbine_blade#Film_cooling film cooling] using some compressor [https://en.wikipedia.org/wiki/Bleed_air bleed air], have to be employed. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seeing what's inside real engines: [https://www.youtube.com/user/AgentJayZ AgentJayZ on youtube]. Huge thanks to you! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines tend to increase efficiency, i.e. higher thrusts for lower fuel consumption, but also try to reduce the exhaust noise. Cost is of course a concern, and an efficiency by itself, but maybe not a hard-constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down with the cost. ===Shaped core or shaped shaft?=== An obvious but important optimization to reduce cost and complexity of manufacturing is to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to design and produce, with a simple [https://en.wikipedia.org/wiki/Lathe lathe] ([https://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [https://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates, for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a shaft turned in raw metal won't change much on its final mass. Moreover, we may use a turbine-level mechanism embedded in the stator to try to cool it, which would make it hollow. The main mechanical issues are probably how to properly fix the blades on rotor and stator, how to fix the rotor on the stator with little gap, and how to balance it/them? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compressor blades. The high-pressure turbine specially have to face very high temperature and pressure. On real engines, they are made of nickel-based [https://en.wikipedia.org/wiki/Superalloys superalloys] or are ceramic-coated. It's the inability of blades to withstand heat and work that limit the power of the engine. Indeed, around 70% of the gas provided by the compressor is used only for chamber and turbine cooling, instead of using it to burn more fuel and create more thrust. The compressor and the turbine are not only made of blades on the rotor, but also blades on the stator. They prevent a rotating air flow driven by the action of rotor blades to form inside the engine, which would decrease the energy of the gas. Stator blades or vanes redirect the airflow on the next stage in the more efficient direction. Highest efficiencies are reached in turbofans when gaps are reduced between rotor blades' tip and the stator, as well as between the stator blades' tip and the rotor. As always, good efficiency means high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. The shape of the blade and the parameters of their cascade also affects the efficiency. A small 5 stage supersonic compressor providing the same pressure rise than a 15 stage subsonic compressor is less efficient, but it may be compensated by the higher thrust-to-weight ratio. Blade geometric design is also very complicated. First turbine engines had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1926, [https://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines Alan A. Griffith] proved that if blades were designed as airfoils, the engine would behave way better, and would even be efficient enough to deserve being built. Airfoils for blade designs allow compressor stages to better increase the static pressure since they create an expander, an increasing area for the air flow to pass through. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C so it will be way off-limits, and even below that, it will deform before melting. And should we mention that blade deformation at high-centrifugal loads, caused by the high rpm, is a very good way to blow the engine off? Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the vanes the turbine, which are, on actual engines, made of a ceramic-coated high-temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge or bleed circulates, enters the blades, and evacuates through small holes in the blades (convective cooling and film cooling). For the rotor blades, the same principle is used, but with compressor air passing inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air or a high speed electric engine for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air and compressor stall, re-ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude. The rotor speed information and altimeter may be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and removed axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we aim. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. See this example of small fan, from [http://www.price-induction.com/site_media/images/dgen-net/technologies/optimisation_de_masse_grand.jpg Price Induction]. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, a simple square-section root is probably enough. Another lead is to create the blade disk and the blades in a single piece. This can be done with modern manufacturing process like electric discharge machining, 5-dof machining or even laser-based 3D printing. Here is an example from Bladon Jets, [http://www.bladonjets.com/technology/blisk/ the BLISK]. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blade-roots are inserted and fixed. Creating a perpendicular blade root is already a challenge. Rotor blades would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting should be considered. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Air bled from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose [[Turbofan:Alternative Designs|alternative turbofan designs]] based on a mixed-flow turbine. Several pages have been created in the [[:Category:Turbofan|Turbofan category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three to five stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while sustaining the combustion. * [[Turbofan:Bearings|Bearings and cooling]]: high speed rotations require adapted bearings and cooling, which may be reused for rotor and even turbine cooling. ===Turbofan design procedure=== # evaluate required thrust (from aircraft mass and lift, but also [[Flight_at_high_altitude|flight characteristics]]) # calculate required mass flow rate for the fan (thust is [https://en.wikipedia.org/wiki/Thrust calculated] from MFR and flow speed) # fix bypass ratio and fan diameter and rpm, thus giving core diameter (BPR may be [[Turbofan:Alternative_Designs#Full_transonic_engine_design_in_a_single_spool_with_2.1_BPR|fixed by design]]) # calculate required power to drive the fan alone # evaluate a gross compressor driving power (refined later)<br />&nbsp; # calculate total power that has to be drawn from the turbine (fan + compressor + losses) # calculate mass flow rate for the combustion alone # calculate mass flow rate for cooling chamber and turbine ## calculate mass flow rate for cooling chamber ## evaluate mass flow rate for cooling turbine to add to the latter ## calculate entropy and fluid parameters at combustor discharge (speed, temperature) ## calculate temperature of turbine vanes and blade and check if it is acceptable ## iterate on item 8.1 until temperature is unacceptable # calculate the number of turbine blades and stages required for this power # calculate compressor discharge pressure and pressure ratio # calculate how many compressor stages are required depending on sonic or supersonic blade design and fix design # calculate compressor driving power # iterate on item 6 until total power varies # design blades for all calculated parameters and re-run at item 6, total power may have changed [[Category:Turbofan]] a8dedbf71228e6f6ceba4786170da0dd76379117 561 560 2015-01-12T00:11:32Z Vincent 1 /* Design versus manufacturing */ minor text updates wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets or other engines. They are however very difficult to design and manufacture and are thus very expensive. On this page, we will explore how costs can be reduced while still having adequate capabilities for high altitude flight. ==General principles== A lot of information is available on Wikipedia's [https://en.wikipedia.org/wiki/Turbofan turbofan page]. General principle is that there is a combustion that puts energy into a gas, this energy is extracted by a turbine, and the turbine drives both the fan that provides thrust and the compression stage that feeds the combustion with oxygen. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake, and so on. The fan provides thrust by creating a massive air flow, and the engine's core also creates thrust by evacuating the high-speed hot combustion gas. In commercial turbofan engines, the fan is generally responsible for 90% of the overall thrust. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [https://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. An engine with a BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost, has the fan mounted on it. One-stage engines exist and are less complicated and expensive to build, but are much less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan, but this is not an easy task due to its very speed. Commercial engines featuring a gearbox for the turbofan's fan ([https://en.wikipedia.org/wiki/Geared_turbofan geared turbofans]) have reached market in the past few years. Multi-spooled engines lessen this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. High-speed fans are complicated to understand and design, and undergo higher losses due to transonic regime, and having a gearbox on a twin-spool engine is even better, to get a high efficiency turbine and a slowly rotating fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas and thus the work that can be extracted by the turbines. * The turbine inlet temperature is the key parameter for turbine engine. Attaining higher temperatures after combustion, in the turbine inlet, makes an engine more efficient, as explained by the [https://en.wikipedia.org/wiki/Brayton_cycle Brayton thermodynamic cycle]. The obvious issue is the temperature that materials can withstand, in particular the first turbine guide vane and the first high-pressure turbine blades. [https://en.wikipedia.org/wiki/High_temperature_metal High-temperature alloys] and special cooling techniques, like [https://en.wikipedia.org/wiki/Turbine_blade#Film_cooling film cooling] using some compressor [https://en.wikipedia.org/wiki/Bleed_air bleed air], have to be employed. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seeing what's inside real engines: [https://www.youtube.com/user/AgentJayZ AgentJayZ on youtube]. Huge thanks to you! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines aim to increase efficiency, reduce noise, life time. Cost is of course a concern, and an efficiency by itself, but not the primary constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down with the cost. ===Shaped core or shaped shaft?=== An obvious but important optimization to reduce cost and complexity of manufacturing is to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to design and produce, with a simple [https://en.wikipedia.org/wiki/Lathe lathe] ([https://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [https://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. These also do not scale down very well. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates, for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. Our small engine design allows us to have a more simple design, since having a shaft turned in raw metal won't change much on its final mass. Moreover, we may use a turbine-level mechanism embedded in the stator to try to cool it, which would make it hollow. The main mechanical issues are probably how to properly fix the blades on rotor and stator with such a small scale and how to fix the rotor on the stator with little gap? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compressor blades. The high-pressure turbine specifically have to face very high temperature and pressure. On real engines, they are made of ceramic-coated nickel-based [https://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine. Indeed, around 70% of the gas provided by the compressor is used only for chamber and turbine cooling, instead of using it to burn more fuel and creating more thrust. The compressor and the turbine are not only made of blades on the rotor, but also blades on the stator, also called vanes. They prevent a rotating air flow driven by the action of rotor blades to form inside the engine, which would not allow pressure to be built on the gas at each stage. Stator blades or vanes redirect the airflow on the next stage in the more efficient direction (see [https://en.wikipedia.org/wiki/Velocity_triangle velocity triangle]). Highest efficiencies are reached in turbofans when gaps are reduced between rotor blades' tip and the stator, as well as between the stator blades' tip and the rotor. As always, good efficiency means high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. The shape of the blades and the parameters of their cascade also affects the efficiency, and in particular the pressure ratios. A small 5 stage supersonic compressor providing the same pressure rise than a 15 stage subsonic compressor is less efficient, but it ''may'' be compensated by the higher thrust-to-weight ratio. Blade geometric design is also very complicated. First turbine engines had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1926, [https://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines Alan A. Griffith] proved that if blades were designed as aerofoils (airfoils in US English), the engine would behave way better, and would even be efficient enough to deserve being built. Airfoils for blade designs allow compressor stages to better increase the static pressure since they create an expander, an increasing area for the air flow to pass through. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C so it will be way off-limits, and even below that, it will deform before melting. And should we mention that blade deformation at high-centrifugal loads, caused by the high rpm, is a very good way to blow the engine off? Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the vanes the turbine, which are, on actual engines, made of a ceramic-coated high-temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge or bleed circulates, enters the blades, and evacuates through small holes in the blades (convective cooling and film cooling). For the rotor blades, the same principle is used, but with compressor air passing inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air or a high speed electric engine for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air and compressor stall, re-ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude. The rotor speed information and altimeter may be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and removed axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we aim. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. See this example of small fan, from [http://www.price-induction.com/site_media/images/dgen-net/technologies/optimisation_de_masse_grand.jpg Price Induction]. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, a simple square-section root is probably enough. Another lead is to create the blade disk and the blades in a single piece. This can be done with modern manufacturing process like electric discharge machining, 5-dof machining or even laser-based 3D printing. Here is an example from Bladon Jets, [http://www.bladonjets.com/technology/blisk/ the BLISK]. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blade-roots are inserted and fixed. Creating a perpendicular blade root is already a challenge. Rotor blades would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting should be considered. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Air bled from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose [[Turbofan:Alternative Designs|alternative turbofan designs]] based on a mixed-flow turbine. Several pages have been created in the [[:Category:Turbofan|Turbofan category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three to five stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while sustaining the combustion. * [[Turbofan:Bearings|Bearings and cooling]]: high speed rotations require adapted bearings and cooling, which may be reused for rotor and even turbine cooling. ===Turbofan design procedure=== # evaluate required thrust (from aircraft mass and lift, but also [[Flight_at_high_altitude|flight characteristics]]) # calculate required mass flow rate for the fan (thust is [https://en.wikipedia.org/wiki/Thrust calculated] from MFR and flow speed) # fix bypass ratio and fan diameter and rpm, thus giving core diameter (BPR may be [[Turbofan:Alternative_Designs#Full_transonic_engine_design_in_a_single_spool_with_2.1_BPR|fixed by design]]) # calculate required power to drive the fan alone # evaluate a gross compressor driving power (refined later)<br />&nbsp; # calculate total power that has to be drawn from the turbine (fan + compressor + losses) # calculate mass flow rate for the combustion alone # calculate mass flow rate for cooling chamber and turbine ## calculate mass flow rate for cooling chamber ## evaluate mass flow rate for cooling turbine to add to the latter ## calculate entropy and fluid parameters at combustor discharge (speed, temperature) ## calculate temperature of turbine vanes and blade and check if it is acceptable ## iterate on item 8.1 until temperature is unacceptable # calculate the number of turbine blades and stages required for this power # calculate compressor discharge pressure and pressure ratio # calculate how many compressor stages are required depending on sonic or supersonic blade design and fix design # calculate compressor driving power # iterate on item 6 until total power varies # design blades for all calculated parameters and re-run at item 6, total power may have changed [[Category:Turbofan]] b7bb112739da93a8da32fce2e1632831de0db99c 562 561 2015-01-12T00:29:59Z Vincent 1 /* Shaped core or shaped shaft? */ hollow shaft for cooling air wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets or other engines. They are however very difficult to design and manufacture and are thus very expensive. On this page, we will explore how costs can be reduced while still having adequate capabilities for high altitude flight. ==General principles== A lot of information is available on Wikipedia's [https://en.wikipedia.org/wiki/Turbofan turbofan page]. General principle is that there is a combustion that puts energy into a gas, this energy is extracted by a turbine, and the turbine drives both the fan that provides thrust and the compression stage that feeds the combustion with oxygen. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake, and so on. The fan provides thrust by creating a massive air flow, and the engine's core also creates thrust by evacuating the high-speed hot combustion gas. In commercial turbofan engines, the fan is generally responsible for 90% of the overall thrust. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [https://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. An engine with a BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost, has the fan mounted on it. One-stage engines exist and are less complicated and expensive to build, but are much less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan, but this is not an easy task due to its very speed. Commercial engines featuring a gearbox for the turbofan's fan ([https://en.wikipedia.org/wiki/Geared_turbofan geared turbofans]) have reached market in the past few years. Multi-spooled engines lessen this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. High-speed fans are complicated to understand and design, and undergo higher losses due to transonic regime, and having a gearbox on a twin-spool engine is even better, to get a high efficiency turbine and a slowly rotating fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas and thus the work that can be extracted by the turbines. * The turbine inlet temperature is the key parameter for turbine engine. Attaining higher temperatures after combustion, in the turbine inlet, makes an engine more efficient, as explained by the [https://en.wikipedia.org/wiki/Brayton_cycle Brayton thermodynamic cycle]. The obvious issue is the temperature that materials can withstand, in particular the first turbine guide vane and the first high-pressure turbine blades. [https://en.wikipedia.org/wiki/High_temperature_metal High-temperature alloys] and special cooling techniques, like [https://en.wikipedia.org/wiki/Turbine_blade#Film_cooling film cooling] using some compressor [https://en.wikipedia.org/wiki/Bleed_air bleed air], have to be employed. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seeing what's inside real engines: [https://www.youtube.com/user/AgentJayZ AgentJayZ on youtube]. Huge thanks to you! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines aim to increase efficiency, reduce noise, life time. Cost is of course a concern, and an efficiency by itself, but not the primary constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down with the cost. ===Shaped core or shaped shaft?=== An obvious but important optimization to reduce cost and complexity of manufacturing is to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to design and produce, with a simple [https://en.wikipedia.org/wiki/Lathe lathe] ([https://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [https://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. These also do not scale down very well. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates, for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. A big advantage of that is that the shaft can pass compressor bleed air to the turbines to use it as coolant. Our small engine design allows us to have a more simple design, since having a shaft turned in raw metal won't change much on its final mass. Moreover, we may use a turbine-level mechanism embedded in the stator to try to cool it, which would make it hollow. The main mechanical issues are probably how to properly fix the blades on rotor and stator with such a small scale and how to fix the rotor on the stator with little gap? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compressor blades. The high-pressure turbine specifically have to face very high temperature and pressure. On real engines, they are made of ceramic-coated nickel-based [https://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine. Indeed, around 70% of the gas provided by the compressor is used only for chamber and turbine cooling, instead of using it to burn more fuel and creating more thrust. The compressor and the turbine are not only made of blades on the rotor, but also blades on the stator, also called vanes. They prevent a rotating air flow driven by the action of rotor blades to form inside the engine, which would not allow pressure to be built on the gas at each stage. Stator blades or vanes redirect the airflow on the next stage in the more efficient direction (see [https://en.wikipedia.org/wiki/Velocity_triangle velocity triangle]). Highest efficiencies are reached in turbofans when gaps are reduced between rotor blades' tip and the stator, as well as between the stator blades' tip and the rotor. As always, good efficiency means high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. The shape of the blades and the parameters of their cascade also affects the efficiency, and in particular the pressure ratios. A small 5 stage supersonic compressor providing the same pressure rise than a 15 stage subsonic compressor is less efficient, but it ''may'' be compensated by the higher thrust-to-weight ratio. Blade geometric design is also very complicated. First turbine engines had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1926, [https://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines Alan A. Griffith] proved that if blades were designed as aerofoils (airfoils in US English), the engine would behave way better, and would even be efficient enough to deserve being built. Airfoils for blade designs allow compressor stages to better increase the static pressure since they create an expander, an increasing area for the air flow to pass through. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C so it will be way off-limits, and even below that, it will deform before melting. And should we mention that blade deformation at high-centrifugal loads, caused by the high rpm, is a very good way to blow the engine off? Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the vanes the turbine, which are, on actual engines, made of a ceramic-coated high-temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge or bleed circulates, enters the blades, and evacuates through small holes in the blades (convective cooling and film cooling). For the rotor blades, the same principle is used, but with compressor air passing inside the rotor. ===Startup and ignition=== Startup can be done at ground manually, with compressed air or a high speed electric engine for example, which will allow to reduce the weight and complexity of the engine. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, and the alternator would be used reversely as a startup DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining, but if pump or throttling malfunction, or more generally if a turbulence in the intake happen, leading to a discontinuous flow of fuel or air and compressor stall, re-ignition would have to be made during the flight. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude. The rotor speed information and altimeter may be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and removed axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we aim. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. See this example of small fan, from [http://www.price-induction.com/site_media/images/dgen-net/technologies/optimisation_de_masse_grand.jpg Price Induction]. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, a simple square-section root is probably enough. Another lead is to create the blade disk and the blades in a single piece. This can be done with modern manufacturing process like electric discharge machining, 5-dof machining or even laser-based 3D printing. Here is an example from Bladon Jets, [http://www.bladonjets.com/technology/blisk/ the BLISK]. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blade-roots are inserted and fixed. Creating a perpendicular blade root is already a challenge. Rotor blades would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting should be considered. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Air bled from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose [[Turbofan:Alternative Designs|alternative turbofan designs]] based on a mixed-flow turbine. Several pages have been created in the [[:Category:Turbofan|Turbofan category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three to five stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while sustaining the combustion. * [[Turbofan:Bearings|Bearings and cooling]]: high speed rotations require adapted bearings and cooling, which may be reused for rotor and even turbine cooling. ===Turbofan design procedure=== # evaluate required thrust (from aircraft mass and lift, but also [[Flight_at_high_altitude|flight characteristics]]) # calculate required mass flow rate for the fan (thust is [https://en.wikipedia.org/wiki/Thrust calculated] from MFR and flow speed) # fix bypass ratio and fan diameter and rpm, thus giving core diameter (BPR may be [[Turbofan:Alternative_Designs#Full_transonic_engine_design_in_a_single_spool_with_2.1_BPR|fixed by design]]) # calculate required power to drive the fan alone # evaluate a gross compressor driving power (refined later)<br />&nbsp; # calculate total power that has to be drawn from the turbine (fan + compressor + losses) # calculate mass flow rate for the combustion alone # calculate mass flow rate for cooling chamber and turbine ## calculate mass flow rate for cooling chamber ## evaluate mass flow rate for cooling turbine to add to the latter ## calculate entropy and fluid parameters at combustor discharge (speed, temperature) ## calculate temperature of turbine vanes and blade and check if it is acceptable ## iterate on item 8.1 until temperature is unacceptable # calculate the number of turbine blades and stages required for this power # calculate compressor discharge pressure and pressure ratio # calculate how many compressor stages are required depending on sonic or supersonic blade design and fix design # calculate compressor driving power # iterate on item 6 until total power varies # design blades for all calculated parameters and re-run at item 6, total power may have changed [[Category:Turbofan]] 7b95f3f584b16c2cbcb1843fe45e677ca9bbf8a5 565 562 2015-01-20T01:08:58Z Vincent 1 /* Startup and ignition */ more details wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets or other engines. They are however very difficult to design and manufacture and are thus very expensive. On this page, we will explore how costs can be reduced while still having adequate capabilities for high altitude flight. ==General principles== A lot of information is available on Wikipedia's [https://en.wikipedia.org/wiki/Turbofan turbofan page]. General principle is that there is a combustion that puts energy into a gas, this energy is extracted by a turbine, and the turbine drives both the fan that provides thrust and the compression stage that feeds the combustion with oxygen. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake, and so on. The fan provides thrust by creating a massive air flow, and the engine's core also creates thrust by evacuating the high-speed hot combustion gas. In commercial turbofan engines, the fan is generally responsible for 90% of the overall thrust. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [https://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. An engine with a BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost, has the fan mounted on it. One-stage engines exist and are less complicated and expensive to build, but are much less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan, but this is not an easy task due to its very speed. Commercial engines featuring a gearbox for the turbofan's fan ([https://en.wikipedia.org/wiki/Geared_turbofan geared turbofans]) have reached market in the past few years. Multi-spooled engines lessen this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. High-speed fans are complicated to understand and design, and undergo higher losses due to transonic regime, and having a gearbox on a twin-spool engine is even better, to get a high efficiency turbine and a slowly rotating fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas and thus the work that can be extracted by the turbines. * The turbine inlet temperature is the key parameter for turbine engine. Attaining higher temperatures after combustion, in the turbine inlet, makes an engine more efficient, as explained by the [https://en.wikipedia.org/wiki/Brayton_cycle Brayton thermodynamic cycle]. The obvious issue is the temperature that materials can withstand, in particular the first turbine guide vane and the first high-pressure turbine blades. [https://en.wikipedia.org/wiki/High_temperature_metal High-temperature alloys] and special cooling techniques, like [https://en.wikipedia.org/wiki/Turbine_blade#Film_cooling film cooling] using some compressor [https://en.wikipedia.org/wiki/Bleed_air bleed air], have to be employed. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seeing what's inside real engines: [https://www.youtube.com/user/AgentJayZ AgentJayZ on youtube]. Huge thanks to you! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines aim to increase efficiency, reduce noise, life time. Cost is of course a concern, and an efficiency by itself, but not the primary constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down with the cost. ===Shaped core or shaped shaft?=== An obvious but important optimization to reduce cost and complexity of manufacturing is to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to design and produce, with a simple [https://en.wikipedia.org/wiki/Lathe lathe] ([https://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [https://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. These also do not scale down very well. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates, for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. A big advantage of that is that the shaft can pass compressor bleed air to the turbines to use it as coolant. Our small engine design allows us to have a more simple design, since having a shaft turned in raw metal won't change much on its final mass. Moreover, we may use a turbine-level mechanism embedded in the stator to try to cool it, which would make it hollow. The main mechanical issues are probably how to properly fix the blades on rotor and stator with such a small scale and how to fix the rotor on the stator with little gap? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compressor blades. The high-pressure turbine specifically have to face very high temperature and pressure. On real engines, they are made of ceramic-coated nickel-based [https://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine. Indeed, around 70% of the gas provided by the compressor is used only for chamber and turbine cooling, instead of using it to burn more fuel and creating more thrust. The compressor and the turbine are not only made of blades on the rotor, but also blades on the stator, also called vanes. They prevent a rotating air flow driven by the action of rotor blades to form inside the engine, which would not allow pressure to be built on the gas at each stage. Stator blades or vanes redirect the airflow on the next stage in the more efficient direction (see [https://en.wikipedia.org/wiki/Velocity_triangle velocity triangle]). Highest efficiencies are reached in turbofans when gaps are reduced between rotor blades' tip and the stator, as well as between the stator blades' tip and the rotor. As always, good efficiency means high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. The shape of the blades and the parameters of their cascade also affects the efficiency, and in particular the pressure ratios. A small 5 stage supersonic compressor providing the same pressure rise than a 15 stage subsonic compressor is less efficient, but it ''may'' be compensated by the higher thrust-to-weight ratio. Blade geometric design is also very complicated. First turbine engines had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1926, [https://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines Alan A. Griffith] proved that if blades were designed as aerofoils (airfoils in US English), the engine would behave way better, and would even be efficient enough to deserve being built. Airfoils for blade designs allow compressor stages to better increase the static pressure since they create an expander, an increasing area for the air flow to pass through. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C so it will be way off-limits, and even below that, it will deform before melting. And should we mention that blade deformation at high-centrifugal loads, caused by the high rpm, is a very good way to blow the engine off? Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the vanes the turbine, which are, on actual engines, made of a ceramic-coated high-temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge or bleed circulates, enters the blades, and evacuates through small holes in the blades (convective cooling and film cooling). For the rotor blades, the same principle is used, but with compressor air passing inside the rotor. ===Start-up and ignition=== Start-up can be done at ground manually, with compressed air or a high speed electric engine for example, which will allow weight and complexity of the engine to be reduced. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, since the alternator would be used reversely as a start-up DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Our main issue for such turbine accessories with a very small engine would be to find the actual space in it where they could fit, especially with a cheap and efficient approach. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining under normal circumstances, but if pump or throttle malfunctions, or if there is a disruption in the intake airflow leading to [https://en.wikipedia.org/wiki/Compressor_stall compressor stall], re-ignition would have to be made during the flight. Free restart of a turbofan is possible if the airspeed is sufficient to make it rotate at the start-up angular velocity, but in other cases that would require an embedded start-up motor or compressed air input. On airliners, bleed air from another engine can be used, but with such tightly constrained motors, we would have a hard time enabling that. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its rotation speed to ensure it's running at an efficient enough value, with regard to altitude (pressure and temperature). That can be done with a rotation sensor, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too. Pressure at different stages would be very useful for engine development, then for behavior indications when running at high altitude. The rotor speed information and altimeter may be redundant with some of the pressure information. ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and removed axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we aim. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. See this example of small fan, from [http://www.price-induction.com/site_media/images/dgen-net/technologies/optimisation_de_masse_grand.jpg Price Induction]. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, a simple square-section root is probably enough. Another lead is to create the blade disk and the blades in a single piece. This can be done with modern manufacturing process like electric discharge machining, 5-dof machining or even laser-based 3D printing. Here is an example from Bladon Jets, [http://www.bladonjets.com/technology/blisk/ the BLISK]. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blade-roots are inserted and fixed. Creating a perpendicular blade root is already a challenge. Rotor blades would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting should be considered. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Air bled from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose [[Turbofan:Alternative Designs|alternative turbofan designs]] based on a mixed-flow turbine. Several pages have been created in the [[:Category:Turbofan|Turbofan category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three to five stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while sustaining the combustion. * [[Turbofan:Bearings|Bearings and cooling]]: high speed rotations require adapted bearings and cooling, which may be reused for rotor and even turbine cooling. ===Turbofan design procedure=== # evaluate required thrust (from aircraft mass and lift, but also [[Flight_at_high_altitude|flight characteristics]]) # calculate required mass flow rate for the fan (thust is [https://en.wikipedia.org/wiki/Thrust calculated] from MFR and flow speed) # fix bypass ratio and fan diameter and rpm, thus giving core diameter (BPR may be [[Turbofan:Alternative_Designs#Full_transonic_engine_design_in_a_single_spool_with_2.1_BPR|fixed by design]]) # calculate required power to drive the fan alone # evaluate a gross compressor driving power (refined later)<br />&nbsp; # calculate total power that has to be drawn from the turbine (fan + compressor + losses) # calculate mass flow rate for the combustion alone # calculate mass flow rate for cooling chamber and turbine ## calculate mass flow rate for cooling chamber ## evaluate mass flow rate for cooling turbine to add to the latter ## calculate entropy and fluid parameters at combustor discharge (speed, temperature) ## calculate temperature of turbine vanes and blade and check if it is acceptable ## iterate on item 8.1 until temperature is unacceptable # calculate the number of turbine blades and stages required for this power # calculate compressor discharge pressure and pressure ratio # calculate how many compressor stages are required depending on sonic or supersonic blade design and fix design # calculate compressor driving power # iterate on item 6 until total power varies # design blades for all calculated parameters and re-run at item 6, total power may have changed [[Category:Turbofan]] c4ce4cf8ca50827c1fa7c374e2b136b814052093 566 565 2015-01-20T01:44:42Z Vincent 1 /* Sensors */ GN&C webcast link wikitext text/x-wiki This page gathers general information on turbofans. Our proposed design is scattered in [[:Category:Turbofan|several pages]], with an index at the [[Build a cheap turbofan#Our design propositions|bottom]] of this page. =How to build a cheap (~ $150) turbofan?= Turbofans are the most efficient engine design for subsonic speeds cruising. They are more powerful and way lighter than reciprocating engines, fly at higher speeds than turbopropellers, and are less fuel-greedy than supersonic-enabled turbojets or other engines. They are however very difficult to design and manufacture and are thus very expensive. On this page, we will explore how costs can be reduced while still having adequate capabilities for high altitude flight. ==General principles== A lot of information is available on Wikipedia's [https://en.wikipedia.org/wiki/Turbofan turbofan page]. General principle is that there is a combustion that puts energy into a gas, this energy is extracted by a turbine, and the turbine drives both the fan that provides thrust and the compression stage that feeds the combustion with oxygen. As air is compressed from the intake, more air becomes available for combustion, and thus create more work on the turbine, and more intake, and so on. The fan provides thrust by creating a massive air flow, and the engine's core also creates thrust by evacuating the high-speed hot combustion gas. In commercial turbofan engines, the fan is generally responsible for 90% of the overall thrust. [[Image:500px-Turbofan_operation.svg.png]] Some design properties and configurations have to be properly calculated depending on the use of the engine, mainly for the intended aircraft speed: * The [https://en.wikipedia.org/wiki/Bypass_ratio Bypass ratio] (BPR) is a ratio between the mass flow rate of air drawn in by the fan but bypassing the engine core to the mass flow rate passing through the engine core. An engine with a BPR = 0 would be a turbojet engine. The higher BPR, the more efficient the engine, but also the slower exhaust speed. * The number of spools: modern engines embed a second and sometimes a third concentric shaft for high pressure operations. The low pressure shaft, the innermost, has the fan mounted on it. One-stage engines exist and are less complicated and expensive to build, but are much less efficient. Indeed, higher rotation speeds in the internal spools allow to provide a more efficient compression. A gearbox may be needed to drive the fan if the shaft has a too important rotation speed in the case of a single-spooled turbofan, but this is not an easy task due to its very speed. Commercial engines featuring a gearbox for the turbofan's fan ([https://en.wikipedia.org/wiki/Geared_turbofan geared turbofans]) have reached market in the past few years. Multi-spooled engines lessen this issue, by keeping the low-pressure stages at relatively low speeds, suited for the fan. High-speed fans are complicated to understand and design, and undergo higher losses due to transonic regime, and having a gearbox on a twin-spool engine is even better, to get a high efficiency turbine and a slowly rotating fan. * The compression ratio is the ratio of the pressure of intake air on compressor discharge air. It is closely determined by the number of stages in the compressor and their efficiency. More compression means more air to blend with fuel and to cool the engine, and even more pressure at output, increasing the speed and mass of output gas and thus the work that can be extracted by the turbines. * The turbine inlet temperature is the key parameter for turbine engine. Attaining higher temperatures after combustion, in the turbine inlet, makes an engine more efficient, as explained by the [https://en.wikipedia.org/wiki/Brayton_cycle Brayton thermodynamic cycle]. The obvious issue is the temperature that materials can withstand, in particular the first turbine guide vane and the first high-pressure turbine blades. [https://en.wikipedia.org/wiki/High_temperature_metal High-temperature alloys] and special cooling techniques, like [https://en.wikipedia.org/wiki/Turbine_blade#Film_cooling film cooling] using some compressor [https://en.wikipedia.org/wiki/Bleed_air bleed air], have to be employed. Turbojet/turbofan engine simulation software from NASA: [http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html EngineSim] A must-read book by Klaus Hünecke: [http://books.google.com/books?id=VpJEm7cFVE4C Jet engines: fundamentals of theory, design, and operation]. Video documentaries from a turbine renovator in Canada, probably the best resource on the Web for seeing what's inside real engines: [https://www.youtube.com/user/AgentJayZ AgentJayZ on youtube]. Huge thanks to you! ==Design versus manufacturing== Design configurations and properties taken into concern on real engines aim to increase efficiency, reduce noise, life time. Cost is of course a concern, and an efficiency by itself, but not the primary constraint as it is for us. Safety of operation is their primary concern, whereas cost and ease of maintenance are our primary concerns -- and maintenance will be an important part of the job if the quality goes down with the cost. ===Shaped core or shaped shaft?=== An obvious but important optimization to reduce cost and complexity of manufacturing is to have a simpler design of the parts creating the gas volume of the engine's core, i.e. the rotor(s) and the stator. In the above schema, we see that the shaft is straight and that the core envelope is curved suit required volume on each stage, although in real life, both are curved. If we take the required volumes on each stage and that we fix the core's envelope shape to a cylinder, the shaft will have a bumped profile (small-large-small diameter). This is much less expensive to design and produce, with a simple [https://en.wikipedia.org/wiki/Lathe lathe] ([https://en.wikipedia.org/wiki/Turning turning]). Earlier engines, like the [https://en.wikipedia.org/wiki/J79 J79], have a cylindrical envelope. A curved envelope is complicated to build, requiring welding, pressing, stage bolting, the same techniques used in stator-construction in modern engines. These also do not scale down very well. Real-world engines don't have a massive turned shaft because of the weight. They consist of plates, for each compressor and turbine stage, that are linked together to the next stage using a cylindrical bolted joint. So basically, the shaft has no core, it's hollow, except for the plates on each stage. A big advantage of that is that the shaft can pass compressor bleed air to the turbines to use it as coolant. Our small engine design allows us to have a more simple design, since having a shaft turned in raw metal won't change much on its final mass. Moreover, we may use a turbine-level mechanism embedded in the stator to try to cool it, which would make it hollow. The main mechanical issues are probably how to properly fix the blades on rotor and stator with such a small scale and how to fix the rotor on the stator with little gap? [[Image:500px-Turbofan_craftedshaft.svg.png]] ===Compressor and turbine blades=== The most complicated parts to build in a turbofan or turbojet engine are the turbine and compressor blades. The high-pressure turbine specifically have to face very high temperature and pressure. On real engines, they are made of ceramic-coated nickel-based [https://en.wikipedia.org/wiki/Superalloys superalloys]. It's the inability of blades to withstand heat and work that limit the power of the engine. Indeed, around 70% of the gas provided by the compressor is used only for chamber and turbine cooling, instead of using it to burn more fuel and creating more thrust. The compressor and the turbine are not only made of blades on the rotor, but also blades on the stator, also called vanes. They prevent a rotating air flow driven by the action of rotor blades to form inside the engine, which would not allow pressure to be built on the gas at each stage. Stator blades or vanes redirect the airflow on the next stage in the more efficient direction (see [https://en.wikipedia.org/wiki/Velocity_triangle velocity triangle]). Highest efficiencies are reached in turbofans when gaps are reduced between rotor blades' tip and the stator, as well as between the stator blades' tip and the rotor. As always, good efficiency means high precision and higher cost. Anyway, the precision of blades will have to be very good if we don't want it to dislocate when it reaches the high rotations-per-minute achieved by such engines. The shape of the blades and the parameters of their cascade also affects the efficiency, and in particular the pressure ratios. A small 5 stage supersonic compressor providing the same pressure rise than a 15 stage subsonic compressor is less efficient, but it ''may'' be compensated by the higher thrust-to-weight ratio. Blade geometric design is also very complicated. First turbine engines had flat blades. At the time, the efficiency of the engine was so terrible that it was believed that turbojets would never beat reciprocating engines. Then, in 1926, [https://en.wikipedia.org/wiki/Alan_Arnold_Griffith#Turbine_engines Alan A. Griffith] proved that if blades were designed as aerofoils (airfoils in US English), the engine would behave way better, and would even be efficient enough to deserve being built. Airfoils for blade designs allow compressor stages to better increase the static pressure since they create an expander, an increasing area for the air flow to pass through. ==Design considerations== ===Temperature control=== Cooling is always needed in turbines, even if recent advances in materials and coatings increased the ability of blades to withstand heat. Since we will use low cost metals, cooling will be the main issue once we figured out how to build the engine. Expected combustion chamber temperature is around 2000°C for hydrocarbon or alcohol fuels. Iron melting point is around 1500°C so it will be way off-limits, and even below that, it will deform before melting. And should we mention that blade deformation at high-centrifugal loads, caused by the high rpm, is a very good way to blow the engine off? Several cooling ways are used in a turbofan/turbojet engine: in the combustion chambers, only a small amount of the actual air flow is used for the combustion, around 20%. The rest is injected on the walls of the chamber and in the end of the combustion to dilute the hot gas, and to prevent the walls from melting (film cooling). Then, the first object struck by this hot gas is the vanes the turbine, which are, on actual engines, made of a ceramic-coated high-temperature alloy, but more importantly, hollow. Blades are welded on the stator ring, around which air from the compressor discharge or bleed circulates, enters the blades, and evacuates through small holes in the blades (convective cooling and film cooling). For the rotor blades, the same principle is used, but with compressor air passing inside the rotor. ===Start-up and ignition=== Start-up can be done at ground manually, with compressed air or a high speed electric engine for example, which will allow weight and complexity of the engine to be reduced. On the other side, a turbine engine is a nice way of having power on-board, using reducing gears and an alternator. That would also reduce the weight required for batteries, since the alternator would be used reversely as a start-up DC motor. Also, the accessories attached to the reduced shaft would allow hydraulic or pneumatic power to be considered. Our main issue for such turbine accessories with a very small engine would be to find the actual space in it where they could fit, especially with a cheap and efficient approach. Igniter mechanisms must be integrated to the engine, possibly a self-maintaining igniter like a thread of tungsten, as used in miniature R/C engines. The combustion should be self-maintaining under normal circumstances, but if pump or throttle malfunctions, or if there is a disruption in the intake airflow leading to [https://en.wikipedia.org/wiki/Compressor_stall compressor stall], re-ignition would have to be made during the flight. Free restart of a turbofan is possible if the airspeed is sufficient to make it rotate at the start-up angular velocity, but in other cases that would require an embedded start-up motor or compressed air input. On airliners, bleed air from another engine can be used, but with such tightly constrained motors, we would have a hard time enabling that. ===Sensors=== Engine must be designed with sensors, at least to determine if the engine is running properly or if it's under failure, and to control its fuel flow rate based on rotation speed, environmental conditions (pressure and temperature) and others. A rotation sensor could a good start, measuring the magnetic field disturbances created by the blades or the rotor. Engine temperature should be controlled and recorded too, probably using a thermocouple. Pressure at different stages would be very useful for engine development and operating, but is also a quality source for failure detection. The rotor speed information and altimeter may be redundant with some of the pressure information. A [http://nescacademy.nasa.gov/video_catalog.php?catid=3&subcatid=1 NASA GN&C webcast] on the Fundamentals of Aircraft Engine Control presents control design, sensors used, and operational safety for turbofan engines ([http://mediaex-server.larc.nasa.gov/Academy/Viewer/?peid=135553bc3b7b4171b7c54ee0578489211d direct link]). ===Fixing blades to rotor=== In real engines, blades are fixed like [http://www.shutterstock.com/pic-9557743/stock-photo-jet-engine.html this], with a dovetail or fir-tree shape that allow them to be mounted and removed axially but not orthogonally. The main problem appearing with this kind of mount is related to the size of the engines we aim. As the diameter of the fan shaft gets smaller, the available space for the blade roots gets smaller, and require a higher precision for their manufacturing. See this example of small fan, from [http://www.price-induction.com/site_media/images/dgen-net/technologies/optimisation_de_masse_grand.jpg Price Induction]. The strength applying to the fixation is luckily reduced due to the small weight of the blades. A simpler design in blade root would be nice for manufacturing ease, a simple square-section root is probably enough. Another lead is to create the blade disk and the blades in a single piece. This can be done with modern manufacturing process like electric discharge machining, 5-dof machining or even laser-based 3D printing. Here is an example from Bladon Jets, [http://www.bladonjets.com/technology/blisk/ the BLISK]. ===Fixing blades to stator=== This is a major issue. On real-size engines, the stator is thick enough to have a rail into which the perpendicular-to-the-blade-roots are inserted and fixed. Creating a perpendicular blade root is already a challenge. Rotor blades would be able to compensate this problem by having longer roots with a locking mechanism on their side, but for the stator, the limited thickness of the stator's wall forbids it. Maybe bolting should be considered. In that case, the screw heads would likely surpass the core's envelope and lightly disturb the fan flow. ===External hardware=== Fuel '''tanks''' in the wings, fuel '''pumps''', fuel '''lines''', and engine '''mounting''' will have to be designed too. Electrical wires for pumps, sensors, ignition and possibly the startup motor/alternator will also be required. Sensors will require input ports on the computer, and pump driving (= engine control loop) will require at least one output port for each engine on the computer. ===Bearings=== Two kinds of bearings are used in turbines. * Ball bearing: stator and rotor are joint using a ball bearing constantly bathed in oil to survive to high speeds/temperature. * Fluid bearing: pressurized oil prevents parts from touching, due to hydrostatic. Longer life and no maintenance, but harder to build and to operate. Carbon or composite lip or blade seals prevent the oil from escaping to other parts of the engine. They may be arranged as labyrinth seals to increase their effect. Air bled from the compressor discharge is often used to counteract the oil pressure on the seals. ==Our design propositions== From the different concerns expressed above, we propose a design for a low-cost turbofan. We also consider and propose [[Turbofan:Alternative Designs|alternative turbofan designs]] based on a mixed-flow turbine. Several pages have been created in the [[:Category:Turbofan|Turbofan category]] to explain each subsystem and parts manufacturability: * [[Turbofan:Compressor|Compressor]]: A three to five stage compressor, with a design allowing easy manufacturing. * [[Turbofan:Blades|Blades]]: How to design an cheaply manufacture compressor, turbine and fan blades. * [[Turbofan:Combustors|Combustors]]: Combustors are the power input of the engine, and need not to melt while sustaining the combustion. * [[Turbofan:Bearings|Bearings and cooling]]: high speed rotations require adapted bearings and cooling, which may be reused for rotor and even turbine cooling. ===Turbofan design procedure=== # evaluate required thrust (from aircraft mass and lift, but also [[Flight_at_high_altitude|flight characteristics]]) # calculate required mass flow rate for the fan (thust is [https://en.wikipedia.org/wiki/Thrust calculated] from MFR and flow speed) # fix bypass ratio and fan diameter and rpm, thus giving core diameter (BPR may be [[Turbofan:Alternative_Designs#Full_transonic_engine_design_in_a_single_spool_with_2.1_BPR|fixed by design]]) # calculate required power to drive the fan alone # evaluate a gross compressor driving power (refined later)<br />&nbsp; # calculate total power that has to be drawn from the turbine (fan + compressor + losses) # calculate mass flow rate for the combustion alone # calculate mass flow rate for cooling chamber and turbine ## calculate mass flow rate for cooling chamber ## evaluate mass flow rate for cooling turbine to add to the latter ## calculate entropy and fluid parameters at combustor discharge (speed, temperature) ## calculate temperature of turbine vanes and blade and check if it is acceptable ## iterate on item 8.1 until temperature is unacceptable # calculate the number of turbine blades and stages required for this power # calculate compressor discharge pressure and pressure ratio # calculate how many compressor stages are required depending on sonic or supersonic blade design and fix design # calculate compressor driving power # iterate on item 6 until total power varies # design blades for all calculated parameters and re-run at item 6, total power may have changed [[Category:Turbofan]] b15494b605ad8a388e1dc0578850d28927a4ee89 0 87 567 2015-02-15T13:30:06Z Vincent 1 Skeleton and Chang'e 3 instruments wikitext text/x-wiki During the first decades of the Moon's exploration, the [https://en.wikipedia.org/wiki/Moon_race Moon race], the largest part of scientific payloads and instruments were used to assess if the lunar environment could support humans and under which conditions. Most instruments were cameras, video of photographic, mapping the surface of the Moon from distance to find suitable landing sites, taking close-up imagery just before crashing on the Moon, or landers taking photographs of the lunar dust. Sample returns were also the most efficient way of studying the lunar regolith. In the past two decades, real science was conducted on the moon or around it, and this page lists the experiments that have taken place since 1950. Soon, the first [http://lunar.xprize.org/ Google Lunax X-Prize] (GLXP) teams shall land on the Moon and will probably increase this list a lot. One of the teams, [http://ptscientists.com/ Part-time Scientists] from Germany, has made a [http://ptscientists.com/go/space call for payloads], cubesat-sized science packages that will land on the moon and return data to earth. The first goal of this page is to have a rough idea of what can be done with minor monetary investment, while still being useful, and possibly answer the call for payload. =Science with direct measurements taken of or from the surface of the Moon= ==From lunar orbit or intersecting trajectory== ''This section is incomplete'' Micrometeroid characterisation and detection by first orbiters. ==From the surface, of the surface== ''This section is incomplete'' ==From the surface, of the rest of the universe== ''This section is incomplete'' In 2013, the Chinese lander [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] had a [https://en.wikipedia.org/wiki/Chang'e_3#Lunar-based_ultraviolet_telescope_.28LUT.29 150mm telescope] onboard and looked at the universe in near ultraviolet wavelengths. It allows for long term monitoring of a target, because the moon has a low rotation rate, and with no UV filtering from Earth atmosphere. In 2013, the Chinese lander [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] had an [https://en.wikipedia.org/wiki/Chang'e_3#Extreme_ultraviolet_.28EUV.29_camera extreme ultraviolet camera] whose purpose was to study the plasmasphere of the Earth and its variation with solar activity. =What we know about the lunar environment= ''This section is incomplete'' [https://en.wikipedia.org/wiki/Topography_of_the_Moon Topography of the Moon] ==Atmosphere== ==Regolith== ==Radiation== ==Magnetic field== ''[https://en.wikipedia.org/wiki/Magnetic_field_of_the_Moon Main article]'' ==Gravity map== 8ef1e2cee49063e7c2ad189d3405e7ad17f6e787 568 567 2015-02-15T17:19:01Z Vincent 1 /* From the surface, of the surface */ spectros wikitext text/x-wiki During the first decades of the Moon's exploration, the [https://en.wikipedia.org/wiki/Moon_race Moon race], the largest part of scientific payloads and instruments were used to assess if the lunar environment could support humans and under which conditions. Most instruments were cameras, video of photographic, mapping the surface of the Moon from distance to find suitable landing sites, taking close-up imagery just before crashing on the Moon, or landers taking photographs of the lunar dust. Sample returns were also the most efficient way of studying the lunar regolith. In the past two decades, real science was conducted on the moon or around it, and this page lists the experiments that have taken place since 1950. Soon, the first [http://lunar.xprize.org/ Google Lunax X-Prize] (GLXP) teams shall land on the Moon and will probably increase this list a lot. One of the teams, [http://ptscientists.com/ Part-time Scientists] from Germany, has made a [http://ptscientists.com/go/space call for payloads], cubesat-sized science packages that will land on the moon and return data to earth. The first goal of this page is to have a rough idea of what can be done with minor monetary investment, while still being useful, and possibly answer the call for payload. =Science with direct measurements taken of or from the surface of the Moon= ==From lunar orbit or intersecting trajectory== ''This section is incomplete'' Micrometeroid characterisation and detection by first orbiters. ==From the surface, of the surface== [https://en.wikipedia.org/wiki/X-ray_spectrometry X-ray spectrometers] provide data very easy to interpret to obtain qualitative information about the elemental composition of a material. Matching data with well-known sample measurements or even standard spectra can lead to quantitative results. The main issue with such devices is producing X-rays, or other sources of energy that will make target elements radiate at X-ray levels of energy. [https://en.wikipedia.org/wiki/Infrared_spectrometry#Practical_IR_spectroscopy Infrared spectrometry] can similarly inform about molecular composition. ''This section is incomplete'' In 2013, the [https://en.wikipedia.org/wiki/Yutu_(rover) YuTu] rover from the Chinese mission [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] carried three instruments: an alpha particle X-ray spectrometer (APXS) described in [http://www.hou.usra.edu/meetings/lpsc2014/pdf/1699.pdf this paper (PDF)] and an infrared spectrometer, both used to analyse the composition of soil or rocks on the surface; a couple of stereo cameras for collision avoidance and trajectory planning. The APXS sensor was the only instrument at the end of the robotic arm. The sensor's head was quite small and had a mass of only 752g. It also comprised a radio-isotope heater unit, to survive the lunar night, adding 390g. The APXS could also be used as a range sensor, measuring the X-ray count rate. The first two spectra can be found in the PDF linked above. ==From the surface, of the rest of the universe== ''This section is incomplete'' In 2013, the Chinese lander [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] had a [https://en.wikipedia.org/wiki/Chang'e_3#Lunar-based_ultraviolet_telescope_.28LUT.29 150mm telescope] onboard and looked at the universe in near ultraviolet wavelengths. It allows for long term monitoring of a target, because the moon has a low rotation rate, and with no UV filtering from Earth atmosphere. In 2013, the Chinese lander [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] had an [https://en.wikipedia.org/wiki/Chang'e_3#Extreme_ultraviolet_.28EUV.29_camera extreme ultraviolet camera] whose purpose was to study the plasmasphere of the Earth and its variation with solar activity. =What we know about the lunar environment= ''This section is incomplete'' [https://en.wikipedia.org/wiki/Topography_of_the_Moon Topography of the Moon] ==Atmosphere== ==Regolith== ==Radiation== ==Magnetic field== ''[https://en.wikipedia.org/wiki/Magnetic_field_of_the_Moon Main article]'' ==Gravity map== 50dfea0a13d0fede461e7783e999f6df99698e5b 569 568 2015-02-15T17:50:45Z Vincent 1 /* From lunar orbit or intersecting trajectory */ chang'e 1 and 2 wikitext text/x-wiki During the first decades of the Moon's exploration, the [https://en.wikipedia.org/wiki/Moon_race Moon race], the largest part of scientific payloads and instruments were used to assess if the lunar environment could support humans and under which conditions. Most instruments were cameras, video of photographic, mapping the surface of the Moon from distance to find suitable landing sites, taking close-up imagery just before crashing on the Moon, or landers taking photographs of the lunar dust. Sample returns were also the most efficient way of studying the lunar regolith. In the past two decades, real science was conducted on the moon or around it, and this page lists the experiments that have taken place since 1950. Soon, the first [http://lunar.xprize.org/ Google Lunax X-Prize] (GLXP) teams shall land on the Moon and will probably increase this list a lot. One of the teams, [http://ptscientists.com/ Part-time Scientists] from Germany, has made a [http://ptscientists.com/go/space call for payloads], cubesat-sized science packages that will land on the moon and return data to earth. The first goal of this page is to have a rough idea of what can be done with minor monetary investment, while still being useful, and possibly answer the call for payload. =Science with direct measurements taken of or from the surface of the Moon= ==From lunar orbit or intersecting trajectory== [https://en.wikipedia.org/wiki/Gamma_ray_spectrometer#Planetary_gamma-ray_spectrometers Gamma-ray spectrometry] can be used from orbit to map the surface distribution of an element. The surface is continuously bombarded by cosmic rays, which make it emit gamma rays, then measured by the spectrometer. ''This section is incomplete'' Micrometeroid characterisation and detection by first orbiters. [https://en.wikipedia.org/wiki/Chang'e_1 Chang'e 1] and 2 had stereo cameras, laser altimeter, imaging spectrometer in visible and near-IR spectrum, Gamma and X-ray spectrometers, microwave radiometer for ground penetration up to 30m, and high energy particle and solar wind detectors. That was the first microwave (radar) mapping of the Moon. ==From the surface, of the surface== [https://en.wikipedia.org/wiki/X-ray_spectrometry X-ray spectrometers] provide data very easy to interpret to obtain qualitative information about the elemental composition of a material. Matching data with well-known sample measurements or even standard spectra can lead to quantitative results. The main issue with such devices is producing X-rays, or other sources of energy that will make target elements radiate at X-ray levels of energy. [https://en.wikipedia.org/wiki/Infrared_spectrometry#Practical_IR_spectroscopy Infrared spectrometry] can similarly inform about molecular composition. ''This section is incomplete'' In 2013, the [https://en.wikipedia.org/wiki/Yutu_(rover) YuTu] rover from the Chinese mission [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] carried three instruments: an alpha particle X-ray spectrometer (APXS) described in [http://www.hou.usra.edu/meetings/lpsc2014/pdf/1699.pdf this paper (PDF)] and an infrared spectrometer, both used to analyse the composition of soil or rocks on the surface; a couple of stereo cameras for collision avoidance and trajectory planning. The APXS sensor was the only instrument at the end of the robotic arm. The sensor's head was quite small and had a mass of only 752g. It also comprised a radio-isotope heater unit, to survive the lunar night, adding 390g. The APXS could also be used as a range sensor, measuring the X-ray count rate. The first two spectra can be found in the PDF linked above. ==From the surface, of the rest of the universe== ''This section is incomplete'' In 2013, the Chinese lander [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] had a [https://en.wikipedia.org/wiki/Chang'e_3#Lunar-based_ultraviolet_telescope_.28LUT.29 150mm telescope] onboard and looked at the universe in near ultraviolet wavelengths. It allows for long term monitoring of a target, because the moon has a low rotation rate, and with no UV filtering from Earth atmosphere. In 2013, the Chinese lander [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] had an [https://en.wikipedia.org/wiki/Chang'e_3#Extreme_ultraviolet_.28EUV.29_camera extreme ultraviolet camera] whose purpose was to study the plasmasphere of the Earth and its variation with solar activity. =What we know about the lunar environment= ''This section is incomplete'' [https://en.wikipedia.org/wiki/Topography_of_the_Moon Topography of the Moon] ==Atmosphere== ==Regolith== ==Radiation== ==Magnetic field== ''[https://en.wikipedia.org/wiki/Magnetic_field_of_the_Moon Main article]'' ==Gravity map== 1f980419f39e5f296024c5c0e2c5ad9150503429 570 569 2015-02-15T20:36:55Z Vincent 1 /* From the surface, of the surface */ link for APXS wikitext text/x-wiki During the first decades of the Moon's exploration, the [https://en.wikipedia.org/wiki/Moon_race Moon race], the largest part of scientific payloads and instruments were used to assess if the lunar environment could support humans and under which conditions. Most instruments were cameras, video of photographic, mapping the surface of the Moon from distance to find suitable landing sites, taking close-up imagery just before crashing on the Moon, or landers taking photographs of the lunar dust. Sample returns were also the most efficient way of studying the lunar regolith. In the past two decades, real science was conducted on the moon or around it, and this page lists the experiments that have taken place since 1950. Soon, the first [http://lunar.xprize.org/ Google Lunax X-Prize] (GLXP) teams shall land on the Moon and will probably increase this list a lot. One of the teams, [http://ptscientists.com/ Part-time Scientists] from Germany, has made a [http://ptscientists.com/go/space call for payloads], cubesat-sized science packages that will land on the moon and return data to earth. The first goal of this page is to have a rough idea of what can be done with minor monetary investment, while still being useful, and possibly answer the call for payload. =Science with direct measurements taken of or from the surface of the Moon= ==From lunar orbit or intersecting trajectory== [https://en.wikipedia.org/wiki/Gamma_ray_spectrometer#Planetary_gamma-ray_spectrometers Gamma-ray spectrometry] can be used from orbit to map the surface distribution of an element. The surface is continuously bombarded by cosmic rays, which make it emit gamma rays, then measured by the spectrometer. ''This section is incomplete'' Micrometeroid characterisation and detection by first orbiters. [https://en.wikipedia.org/wiki/Chang'e_1 Chang'e 1] and 2 had stereo cameras, laser altimeter, imaging spectrometer in visible and near-IR spectrum, Gamma and X-ray spectrometers, microwave radiometer for ground penetration up to 30m, and high energy particle and solar wind detectors. That was the first microwave (radar) mapping of the Moon. ==From the surface, of the surface== [https://en.wikipedia.org/wiki/X-ray_spectrometry X-ray spectrometers] provide data very easy to interpret to obtain qualitative information about the elemental composition of a material. Matching data with well-known sample measurements or even standard spectra can lead to quantitative results. The main issue with such devices is producing X-rays, or other sources of energy that will make target elements radiate at X-ray levels of energy. [https://en.wikipedia.org/wiki/Infrared_spectrometry#Practical_IR_spectroscopy Infrared spectrometry] can similarly inform about molecular composition. ''This section is incomplete'' In 2013, the [https://en.wikipedia.org/wiki/Yutu_(rover) YuTu] rover from the Chinese mission [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] carried three instruments: an [https://en.wikipedia.org/wiki/Alpha_particle_X-ray_spectrometer alpha particle X-ray spectrometer] (APXS) described in [http://www.hou.usra.edu/meetings/lpsc2014/pdf/1699.pdf this paper (PDF)] and an infrared spectrometer, both used to analyse the composition of soil or rocks on the surface; a couple of stereo cameras for collision avoidance and trajectory planning. The APXS sensor was the only instrument at the end of the robotic arm. The sensor's head was quite small and had a mass of only 752g. It also comprised a radio-isotope heater unit, to survive the lunar night, adding 390g. The APXS could also be used as a range sensor, measuring the X-ray count rate. The first two spectra can be found in the PDF linked above. ==From the surface, of the rest of the universe== ''This section is incomplete'' In 2013, the Chinese lander [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] had a [https://en.wikipedia.org/wiki/Chang'e_3#Lunar-based_ultraviolet_telescope_.28LUT.29 150mm telescope] onboard and looked at the universe in near ultraviolet wavelengths. It allows for long term monitoring of a target, because the moon has a low rotation rate, and with no UV filtering from Earth atmosphere. In 2013, the Chinese lander [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] had an [https://en.wikipedia.org/wiki/Chang'e_3#Extreme_ultraviolet_.28EUV.29_camera extreme ultraviolet camera] whose purpose was to study the plasmasphere of the Earth and its variation with solar activity. =What we know about the lunar environment= ''This section is incomplete'' [https://en.wikipedia.org/wiki/Topography_of_the_Moon Topography of the Moon] ==Atmosphere== ==Regolith== ==Radiation== ==Magnetic field== ''[https://en.wikipedia.org/wiki/Magnetic_field_of_the_Moon Main article]'' ==Gravity map== ffaaa8bf3dc499b27582cb69dede07584c03ea64 571 570 2015-02-16T01:18:40Z Vincent 1 /* From the surface, of the rest of the universe */ more chang'e 3 instruments wikitext text/x-wiki During the first decades of the Moon's exploration, the [https://en.wikipedia.org/wiki/Moon_race Moon race], the largest part of scientific payloads and instruments were used to assess if the lunar environment could support humans and under which conditions. Most instruments were cameras, video of photographic, mapping the surface of the Moon from distance to find suitable landing sites, taking close-up imagery just before crashing on the Moon, or landers taking photographs of the lunar dust. Sample returns were also the most efficient way of studying the lunar regolith. In the past two decades, real science was conducted on the moon or around it, and this page lists the experiments that have taken place since 1950. Soon, the first [http://lunar.xprize.org/ Google Lunax X-Prize] (GLXP) teams shall land on the Moon and will probably increase this list a lot. One of the teams, [http://ptscientists.com/ Part-time Scientists] from Germany, has made a [http://ptscientists.com/go/space call for payloads], cubesat-sized science packages that will land on the moon and return data to earth. The first goal of this page is to have a rough idea of what can be done with minor monetary investment, while still being useful, and possibly answer the call for payload. =Science with direct measurements taken of or from the surface of the Moon= ==From lunar orbit or intersecting trajectory== [https://en.wikipedia.org/wiki/Gamma_ray_spectrometer#Planetary_gamma-ray_spectrometers Gamma-ray spectrometry] can be used from orbit to map the surface distribution of an element. The surface is continuously bombarded by cosmic rays, which make it emit gamma rays, then measured by the spectrometer. ''This section is incomplete'' Micrometeroid characterisation and detection by first orbiters. [https://en.wikipedia.org/wiki/Chang'e_1 Chang'e 1] and 2 had stereo cameras, laser altimeter, imaging spectrometer in visible and near-IR spectrum, Gamma and X-ray spectrometers, microwave radiometer for ground penetration up to 30m, and high energy particle and solar wind detectors. That was the first microwave (radar) mapping of the Moon. ==From the surface, of the surface== [https://en.wikipedia.org/wiki/X-ray_spectrometry X-ray spectrometers] provide data very easy to interpret to obtain qualitative information about the elemental composition of a material. Matching data with well-known sample measurements or even standard spectra can lead to quantitative results. The main issue with such devices is producing X-rays, or other sources of energy that will make target elements radiate at X-ray levels of energy. [https://en.wikipedia.org/wiki/Infrared_spectrometry#Practical_IR_spectroscopy Infrared spectrometry] can similarly inform about molecular composition. ''This section is incomplete'' In 2013, the [https://en.wikipedia.org/wiki/Yutu_(rover) YuTu] rover from the Chinese mission [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] carried three instruments: an [https://en.wikipedia.org/wiki/Alpha_particle_X-ray_spectrometer alpha particle X-ray spectrometer] (APXS) described in [http://www.hou.usra.edu/meetings/lpsc2014/pdf/1699.pdf this paper (PDF)] and an infrared spectrometer, both used to analyse the composition of soil or rocks on the surface; a couple of stereo cameras for collision avoidance and trajectory planning. The APXS sensor was the only instrument at the end of the robotic arm. The sensor's head was quite small and had a mass of only 752g. It also comprised a radio-isotope heater unit, to survive the lunar night, adding 390g. The APXS could also be used as a range sensor, measuring the X-ray count rate. The first two spectra can be found in the PDF linked above. ==From the surface, of the rest of the universe== ''This section is incomplete'' '''Lunar Ultraviolet Telescope (LUT)'''. In 1972 Apollo 16 brought a [http://www.lpi.usra.edu/lunar/missions/apollo/apollo_16/experiments/f_ultra/ 76mm far-UV telescope] and in 2013 the [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] lander had a [https://en.wikipedia.org/wiki/Chang'e_3#Lunar-based_ultraviolet_telescope_.28LUT.29 150mm near-UV telescope] onboard. Lunar telescopes allow for long term monitoring of a target, because the moon has a low rotation rate, and with no UV filtering as from Earth atmosphere. Construction details and pictres can be found [http://www.spaceflight101.com/change-3.html here], instruments section. In 2013, the Chinese lander [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] had an [https://en.wikipedia.org/wiki/Chang'e_3#Extreme_ultraviolet_.28EUV.29_camera extreme ultraviolet camera] whose purpose was to study the plasmasphere of the Earth and its variation with solar activity. The camera has a 15 degrees field of view. Details can be found [http://www.spaceflight101.com/change-3.html here], instruments section. =What we know about the lunar environment= ''This section is incomplete'' [https://en.wikipedia.org/wiki/Topography_of_the_Moon Topography of the Moon] ==Atmosphere== ==Regolith== ==Radiation== ==Magnetic field== ''[https://en.wikipedia.org/wiki/Magnetic_field_of_the_Moon Main article]'' ==Gravity map== 29c8f2b023dd3bda432e431ca6dca5ffdb124e2b 572 571 2015-02-16T02:03:01Z Vincent 1 /* From the surface, of the surface */ APXS update wikitext text/x-wiki During the first decades of the Moon's exploration, the [https://en.wikipedia.org/wiki/Moon_race Moon race], the largest part of scientific payloads and instruments were used to assess if the lunar environment could support humans and under which conditions. Most instruments were cameras, video of photographic, mapping the surface of the Moon from distance to find suitable landing sites, taking close-up imagery just before crashing on the Moon, or landers taking photographs of the lunar dust. Sample returns were also the most efficient way of studying the lunar regolith. In the past two decades, real science was conducted on the moon or around it, and this page lists the experiments that have taken place since 1950. Soon, the first [http://lunar.xprize.org/ Google Lunax X-Prize] (GLXP) teams shall land on the Moon and will probably increase this list a lot. One of the teams, [http://ptscientists.com/ Part-time Scientists] from Germany, has made a [http://ptscientists.com/go/space call for payloads], cubesat-sized science packages that will land on the moon and return data to earth. The first goal of this page is to have a rough idea of what can be done with minor monetary investment, while still being useful, and possibly answer the call for payload. =Science with direct measurements taken of or from the surface of the Moon= ==From lunar orbit or intersecting trajectory== [https://en.wikipedia.org/wiki/Gamma_ray_spectrometer#Planetary_gamma-ray_spectrometers Gamma-ray spectrometry] can be used from orbit to map the surface distribution of an element. The surface is continuously bombarded by cosmic rays, which make it emit gamma rays, then measured by the spectrometer. ''This section is incomplete'' Micrometeroid characterisation and detection by first orbiters. [https://en.wikipedia.org/wiki/Chang'e_1 Chang'e 1] and 2 had stereo cameras, laser altimeter, imaging spectrometer in visible and near-IR spectrum, Gamma and X-ray spectrometers, microwave radiometer for ground penetration up to 30m, and high energy particle and solar wind detectors. That was the first microwave (radar) mapping of the Moon. ==From the surface, of the surface== [https://en.wikipedia.org/wiki/X-ray_spectrometry X-ray spectrometers] provide data very easy to interpret to obtain qualitative information about the elemental composition of a material. Matching data with well-known sample measurements or even standard spectra can lead to quantitative results. The main issue with such devices is producing X-rays, or other sources of energy that will make target elements radiate at X-ray levels of energy. [https://en.wikipedia.org/wiki/Infrared_spectrometry#Practical_IR_spectroscopy Infrared spectrometry] can similarly inform about molecular composition. ''This section is incomplete'' In 2013, the [https://en.wikipedia.org/wiki/Yutu_(rover) YuTu] rover from the Chinese mission [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] carried three instruments: an [https://en.wikipedia.org/wiki/Alpha_particle_X-ray_spectrometer alpha particle X-ray spectrometer] (APXS) described in [http://www.hou.usra.edu/meetings/lpsc2014/pdf/1699.pdf this paper (PDF)] and an infrared spectrometer, both used to analyse the composition of soil or rocks on the surface; a couple of stereo cameras for collision avoidance and trajectory planning. The APXS sensor was the only instrument at the end of the robotic arm. The sensor's head was quite small and had a mass of only 752g. It also comprised a radio-isotope heater unit, to survive the lunar night, adding 390g. The APXS could also be used as a range sensor, measuring the X-ray count rate. The first two spectra can be found in the PDF linked above. There was also a part of the instrument in the rover, not in the arm, the total mass is unknown. Details and pictures of the device can be seen [http://www.spaceflight101.com/change-3.html here]. There is a mention of a microwave sonar underneath the YuTu rover [http://www.spaceflight101.com/change-3.html here], but not on the wikipedia page. ==From the surface, of the rest of the universe== ''This section is incomplete'' '''Lunar Ultraviolet Telescope (LUT)'''. In 1972 Apollo 16 brought a [http://www.lpi.usra.edu/lunar/missions/apollo/apollo_16/experiments/f_ultra/ 76mm far-UV telescope] and in 2013 the [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] lander had a [https://en.wikipedia.org/wiki/Chang'e_3#Lunar-based_ultraviolet_telescope_.28LUT.29 150mm near-UV telescope] onboard. Lunar telescopes allow for long term monitoring of a target, because the moon has a low rotation rate, and with no UV filtering as from Earth atmosphere. Construction details and pictres can be found [http://www.spaceflight101.com/change-3.html here], instruments section. In 2013, the Chinese lander [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] had an [https://en.wikipedia.org/wiki/Chang'e_3#Extreme_ultraviolet_.28EUV.29_camera extreme ultraviolet camera] whose purpose was to study the plasmasphere of the Earth and its variation with solar activity. The camera has a 15 degrees field of view. Details can be found [http://www.spaceflight101.com/change-3.html here], instruments section. =What we know about the lunar environment= ''This section is incomplete'' [https://en.wikipedia.org/wiki/Topography_of_the_Moon Topography of the Moon] ==Atmosphere== ==Regolith== ==Radiation== ==Magnetic field== ''[https://en.wikipedia.org/wiki/Magnetic_field_of_the_Moon Main article]'' ==Gravity map== 5b0be176675ccbe01fbc5530d734758f9b161b2a 573 572 2015-03-28T19:35:15Z Vincent 1 /* From lunar orbit or intersecting trajectory */ more from orbit wikitext text/x-wiki During the first decades of the Moon's exploration, the [https://en.wikipedia.org/wiki/Moon_race Moon race], the largest part of scientific payloads and instruments were used to assess if the lunar environment could support humans and under which conditions. Most instruments were cameras, video of photographic, mapping the surface of the Moon from distance to find suitable landing sites, taking close-up imagery just before crashing on the Moon, or landers taking photographs of the lunar dust. Sample returns were also the most efficient way of studying the lunar regolith. In the past two decades, real science was conducted on the moon or around it, and this page lists the experiments that have taken place since 1950. Soon, the first [http://lunar.xprize.org/ Google Lunax X-Prize] (GLXP) teams shall land on the Moon and will probably increase this list a lot. One of the teams, [http://ptscientists.com/ Part-time Scientists] from Germany, has made a [http://ptscientists.com/go/space call for payloads], cubesat-sized science packages that will land on the moon and return data to earth. The first goal of this page is to have a rough idea of what can be done with minor monetary investment, while still being useful, and possibly answer the call for payload. =Science with direct measurements taken of or from the surface of the Moon= ==From lunar orbit or intersecting trajectory== ''This section is incomplete'' [https://en.wikipedia.org/wiki/Gamma_ray_spectrometer#Planetary_gamma-ray_spectrometers Gamma-ray spectrometry] can be used from orbit to map the surface distribution of an element. The surface is continuously bombarded by cosmic rays, which make it emit gamma rays, then measured by the spectrometer. An example use was the mapping of [https://en.wikipedia.org/wiki/KREEP KREEP] distribution by the [https://en.wikipedia.org/wiki/Lunar_Prospector Lunar Prospector] spacecraft. Micrometeroid characterisation and detection by first orbiters. [https://en.wikipedia.org/wiki/Chang'e_1 Chang'e 1] and 2 had stereo cameras, laser altimeter, imaging spectrometer in visible and near-IR spectrum, Gamma and X-ray spectrometers, microwave radiometer for ground penetration up to 30m, and high energy particle and solar wind detectors. That was the first microwave (radar) mapping of the Moon. The electromagnetic environment was also studied from orbit, by [https://en.wikipedia.org/wiki/Chandrayaan-1 Chandrayaan-1], [https://en.wikipedia.org/wiki/Lunar_Prospector Lunar Prospector] and [https://en.wikipedia.org/wiki/LADEE LADEE]. Chandrayaan-1 spacecraft mapped a "mini-magnetosphere" at the Crisium antipode on the moon's far side, using its Sub-keV Atom Reflecting Analyzer (SARA) instrument. Lunar Prospector spacecraft detected changes in the lunar nightside voltage during magnetotail crossings. LADEE has probably done more about magnetic field, to be completed. ==From the surface, of the surface== [https://en.wikipedia.org/wiki/X-ray_spectrometry X-ray spectrometers] provide data very easy to interpret to obtain qualitative information about the elemental composition of a material. Matching data with well-known sample measurements or even standard spectra can lead to quantitative results. The main issue with such devices is producing X-rays, or other sources of energy that will make target elements radiate at X-ray levels of energy. [https://en.wikipedia.org/wiki/Infrared_spectrometry#Practical_IR_spectroscopy Infrared spectrometry] can similarly inform about molecular composition. ''This section is incomplete'' In 2013, the [https://en.wikipedia.org/wiki/Yutu_(rover) YuTu] rover from the Chinese mission [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] carried three instruments: an [https://en.wikipedia.org/wiki/Alpha_particle_X-ray_spectrometer alpha particle X-ray spectrometer] (APXS) described in [http://www.hou.usra.edu/meetings/lpsc2014/pdf/1699.pdf this paper (PDF)] and an infrared spectrometer, both used to analyse the composition of soil or rocks on the surface; a couple of stereo cameras for collision avoidance and trajectory planning. The APXS sensor was the only instrument at the end of the robotic arm. The sensor's head was quite small and had a mass of only 752g. It also comprised a radio-isotope heater unit, to survive the lunar night, adding 390g. The APXS could also be used as a range sensor, measuring the X-ray count rate. The first two spectra can be found in the PDF linked above. There was also a part of the instrument in the rover, not in the arm, the total mass is unknown. Details and pictures of the device can be seen [http://www.spaceflight101.com/change-3.html here]. There is a mention of a microwave sonar underneath the YuTu rover [http://www.spaceflight101.com/change-3.html here], but not on the wikipedia page. ==From the surface, of the rest of the universe== ''This section is incomplete'' '''Lunar Ultraviolet Telescope (LUT)'''. In 1972 Apollo 16 brought a [http://www.lpi.usra.edu/lunar/missions/apollo/apollo_16/experiments/f_ultra/ 76mm far-UV telescope] and in 2013 the [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] lander had a [https://en.wikipedia.org/wiki/Chang'e_3#Lunar-based_ultraviolet_telescope_.28LUT.29 150mm near-UV telescope] onboard. Lunar telescopes allow for long term monitoring of a target, because the moon has a low rotation rate, and with no UV filtering as from Earth atmosphere. Construction details and pictres can be found [http://www.spaceflight101.com/change-3.html here], instruments section. In 2013, the Chinese lander [https://en.wikipedia.org/wiki/Chang'e_3 Chang'e 3] had an [https://en.wikipedia.org/wiki/Chang'e_3#Extreme_ultraviolet_.28EUV.29_camera extreme ultraviolet camera] whose purpose was to study the plasmasphere of the Earth and its variation with solar activity. The camera has a 15 degrees field of view. Details can be found [http://www.spaceflight101.com/change-3.html here], instruments section. =What we know about the lunar environment= ''This section is incomplete'' [https://en.wikipedia.org/wiki/Topography_of_the_Moon Topography of the Moon] ==Atmosphere== ==Regolith== ==Radiation== ==Magnetic field== ''[https://en.wikipedia.org/wiki/Magnetic_field_of_the_Moon Main article]'' ==Gravity map== 29209a34f903f16bd704b98de8989dc006e77b56 6 88 575 2017-11-06T00:45:23Z Vincent 1 This presentation was made at SpaceUp Cote d'Azur (Nice, France) in October 2017. The goal was to explain the challenges of solar system exploration using small satellites, emphasizing that it will cost a lot of money on either launches or reliable spa... wikitext text/x-wiki This presentation was made at SpaceUp Cote d'Azur (Nice, France) in October 2017. The goal was to explain the challenges of solar system exploration using small satellites, emphasizing that it will cost a lot of money on either launches or reliable spacecrafts. 47405e3f4c64dbf9ceee43984686b85135a8b258